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.
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.
NASA Technical Reports Server (NTRS)
Ott, Eric A.
2005-01-01
Scoping of shape changing airfoil concepts including both aerodynamic analysis and materials-related technology assessment effort was performed. Three general categories of potential components were considered-fan blades, booster and compressor blades, and stator airfoils. Based on perceived contributions to improving engine efficiency, the fan blade was chosen as the primary application for a more detailed assessment. A high-level aerodynamic assessment using a GE90-90B Block 4 engine cycle and fan blade geometry indicates that blade camber changes of approximately +/-4deg would be sufficient to result in fan efficiency improvements nearing 1 percent. Constraints related to flight safety and failed mode operation suggest that use of the baseline blade shape with actuation to the optimum cruise condition during a portion of the cycle would be likely required. Application of these conditions to the QAT fan blade and engine cycle was estimated to result in an overall fan efficiency gain of 0.4 percent.
Airfoil shape for flight at subsonic speeds
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.
Airfoil shape for a turbine bucket
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.
Adjoint-based airfoil shape optimization in transonic flow
NASA Astrophysics Data System (ADS)
Gramanzini, Joe-Ray
The primary focus of this work is efficient aerodynamic shape optimization in transonic flow. Adjoint-based optimization techniques are employed on airfoil sections and evaluated in terms of computational accuracy as well as efficiency. This study examines two test cases proposed by the AIAA Aerodynamic Design Optimization Discussion Group. The first is a two-dimensional, transonic, inviscid, non-lifting optimization of a Modified-NACA 0012 airfoil. The second is a two-dimensional, transonic, viscous optimization problem using a RAE 2822 airfoil. The FUN3D CFD code of NASA Langley Research Center is used as the ow solver for the gradient-based optimization cases. Two shape parameterization techniques are employed to study their effect and the number of design variables on the final optimized shape: Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD) and the BandAids free-form deformation technique. For the two airfoil cases, angle of attack is treated as a global design variable. The thickness and camber distributions are the local design variables for MASSOUD, and selected airfoil surface grid points are the local design variables for BandAids. Using the MASSOUD technique, a drag reduction of 72.14% is achieved for the NACA 0012 case, reducing the total number of drag counts from 473.91 to 130.59. Employing the BandAids technique yields a 78.67% drag reduction, from 473.91 to 99.98. The RAE 2822 case exhibited a drag reduction from 217.79 to 132.79 counts, a 39.05% decrease using BandAids.
Airfoil shape for a turbine nozzle
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.
Implicit Shape Parameterization for Kansei Design Methodology
NASA Astrophysics Data System (ADS)
Nordgren, Andreas Kjell; Aoyama, Hideki
Implicit shape parameterization for Kansei design is a procedure that use 3D-models, or concepts, to span a shape space for surfaces in the automotive field. A low-dimensional, yet accurate shape descriptor was found by Principal Component Analysis of an ensemble of point-clouds, which were extracted from mesh-based surfaces modeled in a CAD-program. A theoretical background of the procedure is given along with step-by-step instructions for the required data-processing. The results show that complex surfaces can be described very efficiently, and encode design features by an implicit approach that does not rely on error-prone explicit parameterizations. This provides a very intuitive way to explore shapes for a designer, because various design features can simply be introduced by adding new concepts to the ensemble. Complex shapes have been difficult to analyze with Kansei methods due to the large number of parameters involved, but implicit parameterization of design features provides a low-dimensional shape descriptor for efficient data collection, model-building and analysis of emotional content in 3D-surfaces.
Rapid Parameterization Schemes for Aircraft Shape Optimization
NASA Technical Reports Server (NTRS)
Li, Wu
2012-01-01
A rapid shape parameterization tool called PROTEUS is developed for aircraft shape optimization. This tool can be applied directly to any aircraft geometry that has been defined in PLOT3D format, with the restriction that each aircraft component must be defined by only one data block. PROTEUS has eight types of parameterization schemes: planform, wing surface, twist, body surface, body scaling, body camber line, shifting/scaling, and linear morphing. These parametric schemes can be applied to two types of components: wing-type surfaces (e.g., wing, canard, horizontal tail, vertical tail, and pylon) and body-type surfaces (e.g., fuselage, pod, and nacelle). These schemes permit the easy setup of commonly used shape modification methods, and each customized parametric scheme can be applied to the same type of component for any configuration. This paper explains the mathematics for these parametric schemes and uses two supersonic configurations to demonstrate the application of these schemes.
A Survey of Shape Parameterization Techniques
NASA Technical Reports Server (NTRS)
Samareh, Jamshid A.
1999-01-01
This paper provides a survey of shape parameterization techniques for multidisciplinary optimization and highlights some emerging ideas. The survey focuses on the suitability of available techniques for complex configurations, with suitability criteria based on the efficiency, effectiveness, ease of implementation, and availability of analytical sensitivities for geometry and grids. The paper also contains a section on field grid regeneration, grid deformation, and sensitivity analysis techniques.
Design of a shape adaptive airfoil actuated by a Shape Memory Alloy strip for airplane tail
NASA Astrophysics Data System (ADS)
Shirzadeh, R.; Raissi Charmacani, K.; Tabesh, M.
2011-04-01
Of the factors that mainly affect the efficiency of the wing during a special flow regime, the shape of its airfoil cross section is the most significant. Airfoils are generally designed for a specific flight condition and, therefore, are not fully optimized in all flight conditions. It is very desirable to have an airfoil with the ability to change its shape based on the current regime. Shape memory alloy (SMA) actuators activate in response to changes in the temperature and can recover their original configuration after being deformed. This study presents the development of a method to control the shape of an airfoil using SMA actuators. To predict the thermomechanical behaviors of an SMA thin strip, 3D incremental formulation of the SMA constitutive model is implemented in FEA software package ABAQUS. The interactions between the airfoil structure and SMA thin strip actuator are investigated. Also, the aerodynamic performance of a standard airfoil with a plain flap is compared with an adaptive airfoil.
Two-Dimensional Grids About Airfoils and Other Shapes
NASA Technical Reports Server (NTRS)
Sorenson, R.
1982-01-01
GRAPE computer program generates two-dimensional finite-difference grids about airfoils and other shapes by use of Poisson differential equation. GRAPE can be used with any boundary shape, even one specified by tabulated points and including limited number of sharp corners. Numerically stable and computationally fast, GRAPE provides aerodynamic analyst with efficient and consistant means of grid generation.
General theory of airfoil sections having arbitrary shape or pressure distribution
NASA Technical Reports Server (NTRS)
Allen, H Julian
1945-01-01
In this report a theory of thin airfoils of small camber is developed which permits either the velocity distribution corresponding to a given airfoil shape, or the airfoil shape corresponding to a given velocity distribution to be calculated. The procedures to be employed in these calculations are outlined and illustrated with suitable examples.
Performance Trades Study for Robust Airfoil Shape Optimization
NASA Technical Reports Server (NTRS)
Li, Wu; Padula, Sharon
2003-01-01
From time to time, existing aircraft need to be redesigned for new missions with modified operating conditions such as required lift or cruise speed. This research is motivated by the needs of conceptual and preliminary design teams for smooth airfoil shapes that are similar to the baseline design but have improved drag performance over a range of flight conditions. The proposed modified profile optimization method (MPOM) modifies a large number of design variables to search for nonintuitive performance improvements, while avoiding off-design performance degradation. Given a good initial design, the MPOM generates fairly smooth airfoils that are better than the baseline without making drastic shape changes. Moreover, the MPOM allows users to gain valuable information by exploring performance trades over various design conditions. Four simulation cases of airfoil optimization in transonic viscous ow are included to demonstrate the usefulness of the MPOM as a performance trades study tool. Simulation results are obtained by solving fully turbulent Navier-Stokes equations and the corresponding discrete adjoint equations using an unstructured grid computational fluid dynamics code FUN2D.
Parameterization of 3D brain structures for statistical shape analysis
NASA Astrophysics Data System (ADS)
Zhu, Litao; Jiang, Tianzi
2004-05-01
Statistical Shape Analysis (SSA) is a powerful tool for noninvasive studies of pathophysiology and diagnosis of brain diseases. It also provides a shape constraint for the segmentation of brain structures. There are two key problems in SSA: the representation of shapes and their alignments. The widely used parameterized representations are obtained by preserving angles or areas and the alignments of shapes are achieved by rotating parameter net. However, representations preserving angles or areas do not really guarantee the anatomical correspondence of brain structures. In this paper, we incorporate shape-based landmarks into parameterization of banana-like 3D brain structures to address this problem. Firstly, we get the triangulated surface of the object and extract two landmarks from the mesh, i.e. the ends of the banana-like object. Then the surface is parameterized by creating a continuous and bijective mapping from the surface to a spherical surface based on a heat conduction model. The correspondence of shapes is achieved by mapping the two landmarks to the north and south poles of the sphere and using an extracted origin orientation to select the dateline during parameterization. We apply our approach to the parameterization of lateral ventricle and a multi-resolution shape representation is obtained by using the Discrete Fourier Transform.
Airfoil-shaped micro-mixers for reducing fouling on membrane surfaces
Ho, Clifford K; Altman, Susan J; Clem, Paul G; Hibbs, Michael; Cook, Adam W
2012-10-23
An array of airfoil-shaped micro-mixers that enhances fluid mixing within permeable membrane channels, such as used in reverse-osmosis filtration units, while minimizing additional pressure drop. The enhanced mixing reduces fouling of the membrane surfaces. The airfoil-shaped micro-mixer can also be coated with or comprised of biofouling-resistant (biocidal/germicidal) ingredients.
NASA Technical Reports Server (NTRS)
Derkacs, Thomas (Inventor); Fetheroff, Charles W. (Inventor); Matay, Istvan M. (Inventor); Toth, Istvan J. (Inventor)
1983-01-01
Although the method and apparatus of the present invention can be utilized to apply either a uniform or a nonuniform covering of material over many different workpieces, the apparatus (20) is advantageously utilized to apply a thermal barrier covering (64) to an airfoil (22) which is used in a turbine engine. The airfoil is held by a gripper assembly (86) while a spray gun (24) is effective to apply the covering over the airfoil. When a portion of the covering has been applied, a sensor (28) is utilized to detect the thickness of the covering. A control apparatus (32) compares the thickness of the covering of material which has been applied with the desired thickness and is subsequently effective to regulate the operation of the spray gun to adaptively apply a covering of a desired thickness with an accuracy of at least plus or minus 0.0015 of an inch (1.5 mils) despite unanticipated process variations.
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.
Family of airfoil shapes for rotating blades. [for increased power efficiency and blade stability
NASA Technical Reports Server (NTRS)
Noonan, K. W. (Inventor)
1983-01-01
An airfoil which has particular application to the blade or blades of rotor aircraft such as helicopters and aircraft propellers is described. The airfoil thickness distribution and camber are shaped to maintain a near zero pitching moment coefficient over a wide range of lift coefficients and provide a zero pitching moment coefficient at section Mach numbers near 0.80 and to increase the drag divergence Mach number resulting in superior aircraft performance.
Numerical computation of viscous flow about unconventional airfoil shapes
NASA Technical Reports Server (NTRS)
Ahmed, S.; Tannehill, J. C.
1990-01-01
A new two-dimensional computer code was developed to analyze the viscous flow around unconventional airfoils at various Mach numbers and angles of attack. The Navier-Stokes equations are solved using an implicit, upwind, finite-volume scheme. Both laminar and turbulent flows can be computed. A new nonequilibrium turbulence closure model was developed for computing turbulent flows. This two-layer eddy viscosity model was motivated by the success of the Johnson-King model in separated flow regions. The influence of history effects are described by an ordinary differential equation developed from the turbulent kinetic energy equation. The performance of the present code was evaluated by solving the flow around three airfoils using the Reynolds time-averaged Navier-Stokes equations. Excellent results were obtained for both attached and separated flows about the NACA 0012 airfoil, the RAE 2822 airfoil, and the Integrated Technology A 153W airfoil. Based on the comparison of the numerical solutions with the available experimental data, it is concluded that the present code in conjunction with the new nonequilibrium turbulence model gives excellent results.
The leading-edge stall of airfoils with various nose shapes
NASA Astrophysics Data System (ADS)
Kraljic, Matthew; Rusak, Zvi; Wang, Shixiao
2015-11-01
We study the inception of leading-edge stall on stationary, smooth thin airfoils with various nose shapes of the form xa (where 0 < a < 1 / 2) at low to moderately high chord Reynolds number flows. A reduced-order, multi-scale model problem is developed and solved using numerical simulations. The asymptotic theory demonstrates that a subsonic flow about a thin airfoil can be described in terms of an outer region, around most of the airfoil's chord, and an inner region, around the nose, that asymptotically match each other. The flow in the outer region is dominated by the classical thin airfoil theory. Scaled (magnified) coordinates and a modified (smaller) Reynolds number ReM are used to correctly account for the nonlinear behavior and extreme velocity changes in the inner region, where both the near-stagnation and high suction areas occur. The inner region problem is solved numerically to determine the inception of leading-edge stall on the nose. It is found that stall is delayed to higher angles of attack with the decrease of nose parameter a. Specifically, new airfoil shapes are proposed with increased stall angle at subsonic speeds and higher critical Mach numbers at transonic speeds.
Estimation of morphing airfoil shape and aerodynamic load using artificial hair sensors
NASA Astrophysics Data System (ADS)
Butler, Nathan S.; Su, Weihua; Thapa Magar, Kaman S.; Reich, Gregory W.
2016-04-01
An active area of research in adaptive structures focuses on the use of continuous wing shape changing methods as a means of replacing conventional discrete control surfaces and increasing aerodynamic efficiency. Although many shape-changing methods have been used since the beginning of heavier-than-air flight, the concept of performing camber actuation on a fully-deformable airfoil has not been widely applied. A fundamental problem of applying this concept to real-world scenarios is the fact that camber actuation is a continuous, time-dependent process. Therefore, if camber actuation is to be used in a closed-loop feedback system, one must be able to determine the instantaneous airfoil shape as well as the aerodynamic loads at all times. One approach is to utilize a new type of artificial hair sensors developed at the Air Force Research Laboratory to determine the flow conditions surrounding deformable airfoils. In this work, the hair sensor measurement data will be simulated by using the flow solver XFoil, with the assumption that perfect data with no noise can be collected from the hair sensor measurements. Such measurements will then be used in an artificial neural network based process to approximate the instantaneous airfoil camber shape, lift coefficient, and moment coefficient at a given angle of attack. Various aerodynamic and geometrical properties approximated from the artificial hair sensor and artificial neural network system will be compared with the results of XFoil in order to validate the approximation approach.
A Feasibility Study to Control Airfoil Shape Using THUNDER
NASA Technical Reports Server (NTRS)
Pinkerton, Jennifer L.; Moses, Robert W.
1997-01-01
The objective of this study was to assess the capabilities of a new out-of-plane displacement piezoelectric actuator called thin-layer composite-unimorph ferroelectric driver and sensor (THUNDER) to alter the upper surface geometry of a subscale airfoil to enhance performance under aerodynamic loading. Sixty test conditions, consisting of combinations of five angles of attack, four dc applied voltages, and three tunnel velocities, were studied in a tabletop wind tunnel. Results indicated that larger magnitudes of applied voltage produced larger wafer displacements. Wind-off displacements were also consistently larger than wind-on. Higher velocities produced larger displacements than lower velocities because of increased upper surface suction. Increased suction also resulted in larger displacements at higher angles of attack. Creep and hysteresis of the wafer, which were identified at each test condition, contributed to larger negative displacements for all negative applied voltages and larger positive displacements for the smaller positive applied voltage (+102 V). An elastic membrane used to hold the wafer to the upper surface hindered displacements at the larger positive applied voltage (+170 V). Both creep and hysteresis appeared bounded based on the analysis of several displacement cycles. These results show that THUNDER can be used to alter the camber of a small airfoil under aerodynamic loads.
NASA Astrophysics Data System (ADS)
Zhang, Qiang
The effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape on the aerodynamic performance of turbine airfoils are investigated in compressible, high speed flows. The University of Utah Transonic Wind Tunnel is employed for the experimental part of the study. Two different test sections are designed to produce Mach numbers, Reynolds numbers, passage mass flow rates, and physical dimensions, which match values along turbine blades in operating engines: (i) a nonturning test section with a symmetric airfoil, and (ii) a cascade test section with a cambered turbine vane. The nonuniform, irregular, three-dimensional surface roughness is characterized using the equivalent sand grain roughness size. Changing the airfoil surface roughness condition has a substantial effect on wake profiles of total pressure loss coefficients, normalized Mach number, normalized kinetic energy, and on the normalized and dimensional magnitudes of Integrated Aerodynamic Losses produced by the airfoils. Comparisons with results for a symmetric airfoil and a cambered vane show that roughness has more substantial effects on losses produced by the symmetric airfoil than the cambered vane. Data are also provided that illustrate the larger loss magnitudes are generally present with flow turning and cambered airfoils, than with symmetric airfoils. Wake turbulence structure of symmetric airfoils and cambered vanes are also studied experimentally. The effects of surface roughness and freestream turbulence levels on wake distributions of mean velocity, turbulence intensity, and power spectral density profiles and vortex shedding frequencies are quantified one axial chord length downstream of the test airfoils. As the level of surface roughness increases, all wake profile quantities broaden significantly and nondimensional vortex shedding frequencies decrease. Wake profiles produced by the symmetric airfoil are more sensitive to variations of surface
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.
Aerodynamic effects of simulated ice shapes on two-dimensional airfoils and a swept finite tail
NASA Astrophysics Data System (ADS)
Alansatan, Sait
An experimental study was conducted to investigate the effect of simulated glaze ice shapes on the aerodynamic performance characteristics of two-dimensional airfoils and a swept finite tail. The two dimensional tests involved two NACA 0011 airfoils with chords of 24 and 12 inches. Glaze ice shapes computed with the LEWICE code that were representative of 22.5-min and 45-min ice accretions were simulated with spoilers, which were sized to approximate the horn heights of the LEWICE ice shapes. Lift, drag, pitching moment, and surface pressure coefficients were obtained for a range of test conditions. Test variables included Reynolds number, geometric scaling, control deflection and the key glaze ice features, which were horn height, horn angle, and horn location. For the three-dimensional tests, a 25%-scale business jet empennage (BJE) with a T-tail configuration was used to study the effect of ice shapes on the aerodynamic performance of a swept horizontal tail. Simulated glaze ice shapes included the LEWICE and spoiler ice shapes to represent 9-min and 22.5-min ice accretions. Additional test variables included Reynolds number and elevator deflection. Lift, drag, hinge moment coefficients as well as boundary layer velocity profiles were obtained. The experimental results showed substantial degradation in aerodynamic performance of the airfoils and the swept horizontal tail due to the simulated ice shapes. For the two-dimensional airfoils, the largest aerodynamic penalties were obtained when the 3-in spoiler-ice, which was representative of 45-min glaze ice accretions, was set normal to the chord. Scale and Reynolds effects were not significant for lift and drag. However, pitching moments and pressure distributions showed great sensitivity to Reynolds number and geometric scaling. For the threedimensional study with the swept finite tail, the 22.5-min ice shapes resulted in greater aerodynamic performance degradation than the 9-min ice shapes. The addition of 24
Computer-aided roll pass design in rolling of airfoil shapes
NASA Technical Reports Server (NTRS)
Akgerman, N.; Lahoti, G. D.; Altan, T.
1980-01-01
This paper describes two computer-aided design (CAD) programs developed for modeling the shape rolling process for airfoil sections. The first program, SHPROL, uses a modular upper-bound method of analysis and predicts the lateral spread, elongation, and roll torque. The second program, ROLPAS, predicts the stresses, roll separating force, the roll torque and the details of metal flow by simulating the rolling process, using the slab method of analysis. ROLPAS is an interactive program; it offers graphic display capabilities and allows the user to interact with the computer via a keyboard, CRT, and a light pen. The accuracy of the computerized models was evaluated by (a) rolling a selected airfoil shape at room temperature from 1018 steel and isothermally at high temperature from Ti-6Al-4V, and (b) comparing the experimental results with computer predictions. The comparisons indicated that the CAD systems, described here, are useful for practical engineering purposes and can be utilized in roll pass design and analysis for airfoil and similar shapes.
Illustration of airfoil shape effect on forward-swept wing divergence
NASA Technical Reports Server (NTRS)
Bland, S. R.
1980-01-01
A static aeroelastic analysis is presented of the divergence of untapered wings with conventional and supercritical airfoil sections at sweep angles of zero and -15 deg. One bending and one torsion mode were employed for a uniform rectangular cantilevered beam with the elastic axis at midchord, and calculations were based on a two-dimensional differential equations formulation in the structural coordinate system and in simple strip theory. A minimum divergence speed in the transonic range is obtained which is associated with the rearward shift of the aerodynamic center, and a 17% difference in minimum divergence dynamic pressure is found between a supercritical and a conventional wing. It is noted that although the strip method employed allows the assessment of the sensitivity of airfoil shapes to divergence, three-dimensional transonic aerodynamic methods should be used to predict wing divergence characteristics.
Controlled Aeroelastic Response and Airfoil Shaping Using Adaptive Materials and Integrated Systems
NASA Technical Reports Server (NTRS)
Pinkerton, Jennifer L.; McGowan, Anna-Maria R.; Moses, Robert W.; Scott, Robert C.; Heeg, Jennifer
1996-01-01
This paper presents an overview of several activities of the Aeroelasticity Branch at the NASA Langley Research Center in the area of applying adaptive materials and integrated systems for controlling both aircraft aeroelastic response and airfoil shape. The experimental results of four programs are discussed: the Piezoelectric Aeroelastic Response Tailoring Investigation (PARTI); the Adaptive Neural Control of Aeroelastic Response (ANCAR) program; the Actively Controlled Response of Buffet Affected Tails (ACROBAT) program; and the Airfoil THUNDER Testing to Ascertain Characteristics (ATTACH) project. The PARTI program demonstrated active flutter control and significant rcductions in aeroelastic response at dynamic pressures below flutter using piezoelectric actuators. The ANCAR program seeks to demonstrate the effectiveness of using neural networks to schedule flutter suppression control laws. Th,e ACROBAT program studied the effectiveness of a number of candidate actuators, including a rudder and piezoelectric actuators, to alleviate vertical tail buffeting. In the ATTACH project, the feasibility of using Thin-Layer Composite-Uimorph Piezoelectric Driver and Sensor (THUNDER) wafers to control airfoil aerodynamic characteristics was investigated. Plans for future applications are also discussed.
Wind tunnel evaluation of air-foil performance using simulated ice shapes
NASA Technical Reports Server (NTRS)
Bragg, M. B.; Zaguli, R. J.; Gregorek, G. M.
1982-01-01
A two-phase wind tunnel test was conducted in the 6 by 9 foot Icing Research Tunnel (IRT) at NASA Lewis Research Center to evaluate the effect of ice on the performance of a full scale general aviation wing. In the first IRT tests, rime and glaze shapes were carefully documented as functions of angle of attack and free stream conditions. Next, simulated ice shapes were constructed for two rime and two glaze shapes and used in the second IRT tunnel entry. The ice shapes and the clean airfoil were tapped to obtain surface pressures and a probe used to measure the wake characteristics. These data were recorded and processed, on-line, with a minicomputer/digital data acquisition system. The effect of both rime and glaze ice on the pressure distribution, Cl, Cd, and Cm are presented.
NASA Technical Reports Server (NTRS)
Lahoti, G. D.; Akgerman, N.; Altan, T.
1978-01-01
Mild steel (AISI 1018) was selected as model cold-rolling material and Ti-6Al-4V and INCONEL 718 were selected as typical hot-rolling and cold-rolling alloys, respectively. The flow stress and workability of these alloys were characterized and friction factor at the roll/workpiece interface was determined at their respective working conditions by conducting ring tests. Computer-aided mathematical models for predicting metal flow and stresses, and for simulating the shape-rolling process were developed. These models utilize the upper-bound and the slab methods of analysis, and are capable of predicting the lateral spread, roll-separating force, roll torque and local stresses, strains and strain rates. This computer-aided design (CAD) system is also capable of simulating the actual rolling process and thereby designing roll-pass schedule in rolling of an airfoil or similar shape. The predictions from the CAD system were verified with respect to cold rolling of mild steel plates. The system is being applied to cold and hot isothermal rolling of an airfoil shape, and will be verified with respect to laboratory experiments under controlled conditions.
Ice shapes and the resulting drag increase for a NACA 0012 airfoil
NASA Technical Reports Server (NTRS)
Olsen, W.; Shaw, R. J.; Newton, J.
1984-01-01
Experimental measurements of the ice shapes and resulting drag increases were measured in the NASA-Lewis Icing Research Tunnel. The measurements were made over a large range of conditions (e.g., airspeed and temperature, drop size and liquid water content of the cloud, and the angle of attack of the airfoil). The measured drag increase did not agree with the existing correlation. Additional results were given which are helpful in understanding the ice structure and the way it forms, and in improving the ice accretion modeling theories. There are data on the ice surface roughness, on the effect of the ice shape on the local droplet catch, and on the relative importance of various parts of the ice shape on the drag increase. Experimental repeatability is also discussed.
NASA Technical Reports Server (NTRS)
Lahoti, G. D.; Akgerman, N.; Altan, T.
1978-01-01
Mild steel (AISI 1018) was selected as model cold rolling material and Ti-6A1-4V and Inconel 718 were selected as typical hot rolling and cold rolling alloys, respectively. The flow stress and workability of these alloys were characterized and friction factor at the roll/workpiece interface was determined at their respective working conditions by conducting ring tests. Computer-aided mathematical models for predicting metal flow and stresses, and for simulating the shape rolling process were developed. These models utilized the upper bound and the slab methods of analysis, and were capable of predicting the lateral spread, roll separating force, roll torque, and local stresses, strains and strain rates. This computer-aided design system was also capable of simulating the actual rolling process, and thereby designing the roll pass schedule in rolling of an airfoil or a similar shape.
GRAPE- TWO-DIMENSIONAL GRIDS ABOUT AIRFOILS AND OTHER SHAPES BY THE USE OF POISSON'S EQUATION
NASA Technical Reports Server (NTRS)
Sorenson, R. L.
1994-01-01
The ability to treat arbitrary boundary shapes is one of the most desirable characteristics of a method for generating grids, including those about airfoils. In a grid used for computing aerodynamic flow over an airfoil, or any other body shape, the surface of the body is usually treated as an inner boundary and often cannot be easily represented as an analytic function. The GRAPE computer program was developed to incorporate a method for generating two-dimensional finite-difference grids about airfoils and other shapes by the use of the Poisson differential equation. GRAPE can be used with any boundary shape, even one specified by tabulated points and including a limited number of sharp corners. The GRAPE program has been developed to be numerically stable and computationally fast. GRAPE can provide the aerodynamic analyst with an efficient and consistent means of grid generation. The GRAPE procedure generates a grid between an inner and an outer boundary by utilizing an iterative procedure to solve the Poisson differential equation subject to geometrical restraints. In this method, the inhomogeneous terms of the equation are automatically chosen such that two important effects are imposed on the grid. The first effect is control of the spacing between mesh points along mesh lines intersecting the boundaries. The second effect is control of the angles with which mesh lines intersect the boundaries. Along with the iterative solution to Poisson's equation, a technique of coarse-fine sequencing is employed to accelerate numerical convergence. GRAPE program control cards and input data are entered via the NAMELIST feature. Each variable has a default value such that user supplied data is kept to a minimum. Basic input data consists of the boundary specification, mesh point spacings on the boundaries, and mesh line angles at the boundaries. Output consists of a dataset containing the grid data and, if requested, a plot of the generated mesh. The GRAPE program is
NASA Technical Reports Server (NTRS)
Kaattari, G. E.
1973-01-01
A method is presented for determining shock envelopes and pressure distributions for two-dimensional airfoils at angles of attack sufficiently large to cause shock detachment and subsonic flow over the windward surface of the airfoil. Correlation functions obtained from exact solutions are used to relate the shock standoff distance at the stagnation and sonic points of the body through a suitable choice for the shock shape. The necessary correlation functions were obtained from perfect gas solutions but may be extended to any gas flow for which the normal shock-density ratio can be specified.
NASA Technical Reports Server (NTRS)
Knight, Montgomery; Harris, Thomas A
1931-01-01
This experimental investigation was conducted primarily for the purpose of obtaining a method of correcting to free air conditions the results of airfoil force tests in four open wind tunnel jets of different shapes. Tests were also made to determine whether the jet boundaries had any appreciable effect on the pitching moments of a complete airplane model. Satisfactory corrections for the effect of the boundaries of the various jets were obtained for all the airfoils tested, the span of the largest being 0.75 of the jet width. The corrections for angle of attack were, in general, larger than those for drag. The boundaries had no appreciable effect on the pitching moments of either the airfoils or the complete airplane model. Increasing turbulence appeared to increase the minimum drag and maximum lift and to decrease the pitching moment.
Computational Aerodynamic Analysis of Three-Dimensional Ice Shapes on a NACA 23012 Airfoil
NASA Technical Reports Server (NTRS)
Jun, Garam; Oliden, Daniel; Potapczuk, Mark G.; Tsao, Jen-Ching
2014-01-01
The present study identifies a process for performing computational fluid dynamic calculations of the flow over full three-dimensional (3D) representations of complex ice shapes deposited on aircraft surfaces. Rime and glaze icing geometries formed on a NACA23012 airfoil were obtained during testing in the NASA Glenn Research Center's Icing Research Tunnel (IRT). The ice shape geometries were scanned as a cloud of data points using a 3D laser scanner. The data point clouds were meshed using Geomagic software to create highly accurate models of the ice surface. The surface data was imported into Pointwise grid generation software to create the CFD surface and volume grids. It was determined that generating grids in Pointwise for complex 3D icing geometries was possible using various techniques that depended on the ice shape. Computations of the flow fields over these ice shapes were performed using the NASA National Combustion Code (NCC). Results for a rime ice shape for angle of attack conditions ranging from 0 to 10 degrees and for freestream Mach numbers of 0.10 and 0.18 are presented. For validation of the computational results, comparisons were made to test results from rapid-prototype models of the selected ice accretion shapes, obtained from a separate study in a subsonic wind tunnel at the University of Illinois at Urbana-Champaign. The computational and experimental results were compared for values of pressure coefficient and lift. Initial results show fairly good agreement for rime ice accretion simulations across the range of conditions examined. The glaze ice results are promising but require some further examination.
Computational Aerodynamic Analysis of Three-Dimensional Ice Shapes on a NACA 23012 Airfoil
NASA Technical Reports Server (NTRS)
Jun, GaRam; Oliden, Daniel; Potapczuk, Mark G.; Tsao, Jen-Ching
2014-01-01
The present study identifies a process for performing computational fluid dynamic calculations of the flow over full three-dimensional (3D) representations of complex ice shapes deposited on aircraft surfaces. Rime and glaze icing geometries formed on a NACA23012 airfoil were obtained during testing in the NASA Glenn Research Centers Icing Research Tunnel (IRT). The ice shape geometries were scanned as a cloud of data points using a 3D laser scanner. The data point clouds were meshed using Geomagic software to create highly accurate models of the ice surface. The surface data was imported into Pointwise grid generation software to create the CFD surface and volume grids. It was determined that generating grids in Pointwise for complex 3D icing geometries was possible using various techniques that depended on the ice shape. Computations of the flow fields over these ice shapes were performed using the NASA National Combustion Code (NCC). Results for a rime ice shape for angle of attack conditions ranging from 0 to 10 degrees and for freestream Mach numbers of 0.10 and 0.18 are presented. For validation of the computational results, comparisons were made to test results from rapid-prototype models of the selected ice accretion shapes, obtained from a separate study in a subsonic wind tunnel at the University of Illinois at Urbana-Champaign. The computational and experimental results were compared for values of pressure coefficient and lift. Initial results show fairly good agreement for rime ice accretion simulations across the range of conditions examined. The glaze ice results are promising but require some further examination.
Airfoil-Shaped Fluid Flow Tool for Use in Making Differential Measurements
NASA Technical Reports Server (NTRS)
England, John Dwight (Inventor); Kelley, Anthony R. (Inventor); Cronise, Raymond J. (Inventor)
2014-01-01
A fluid flow tool includes an airfoil structure and a support arm. The airfoil structure's high-pressure side and low-pressure side are positioned in a conduit by the support arm coupled to the conduit. The high-pressure and low-pressure sides substantially face opposing walls of the conduit. At least one measurement port is formed in the airfoil structure at each of its high-pressure side and low-pressure side. A first manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the high-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit. A second manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the low-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit.
NASA Technical Reports Server (NTRS)
Papadakis, M.; Breer, M.; Craig, N.; Liu, X.
1994-01-01
An experimental method has been developed to determine the water droplet impingement characteristics on two- and three-dimensional aircraft surfaces. The experimental water droplet impingement data are used to validate particle trajectory analysis codes that are used in aircraft icing analyses and engine inlet particle separator analyses. The aircraft surface is covered with thin strips of blotter paper in areas of interest. The surface is then exposed to an airstream that contains a dyed-water spray cloud. The water droplet impingement data are extracted from the dyed blotter paper strips by measuring the optical reflectance of each strip with an automated reflectometer. Experimental impingement efficiency data represented for a NLF (1)-0414 airfoil, a swept MS (1)-0317 airfoil, a Boeing 737-300 engine inlet model, two simulated ice shapes and a swept NACA 0012 wingtip. Analytical impingement efficiency data are also presented for the NLF (1)-0414 airfoil and the Boeing 737-300 engine inlet model.
Finding optimum airfoil shape to get maximum aerodynamic efficiency for a wind turbine
NASA Astrophysics Data System (ADS)
Sogukpinar, Haci; Bozkurt, Ismail
2017-02-01
In this study, aerodynamic performances of S-series wind turbine airfoil of S 825 are investigated to find optimum angle of attack. Aerodynamic performances calculations are carried out by utilization of a Computational Fluid Dynamics (CFD) method withstand finite capacity approximation by using Reynolds-Averaged-Navier Stokes (RANS) theorem. The lift and pressure coefficients, lift to drag ratio of airfoil S 825 are analyzed with SST turbulence model then obtained results crosscheck with wind tunnel data to verify the precision of computational Fluid Dynamics (CFD) approximation. The comparison indicates that SST turbulence model used in this study can predict aerodynamics properties of wind blade.
Analytical Investigation of Icing Limit for Diamond-Shaped Airfoil in Transonic and Supersonic Flow
NASA Technical Reports Server (NTRS)
Callaghan, Edmund E.; Serafini, John S.
1953-01-01
Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.
NASA Technical Reports Server (NTRS)
Sorenson, R. L.
1980-01-01
A method for generating two dimensional finite difference grids about airfoils and other shapes by the use of the Poisson differential equation is developed. The inhomogeneous terms are automatically chosen such that two important effects are imposed on the grid at both the inner and outer boundaries. The first effect is control of the spacing between mesh points along mesh lines intersecting the boundaries. The second effect is control of the angles with which mesh lines intersect the boundaries. A FORTRAN computer program has been written to use this method. A description of the program, a discussion of the control parameters, and a set of sample cases are included.
Wu, Linhui; Wan, Wenbo; Wang, Xin; Zhou, Zhongxing; Li, Jiao; Zhang, Limin; Zhao, Huijuan; Gao, Feng
2014-01-01
A fundamental approach to enhancing the sensitivity of the fluorescence molecular tomography (FMT) is to incorporate diffuse optical tomography (DOT) to modify the light propagation modeling. However, the traditional voxel-based DOT has been involving a severely ill-posed inverse problem and cannot retrieve the optical property distributions with the acceptable quantitative accuracy and spatial resolution. Although, with the aid of an anatomical imaging modality, the structural-prior-based DOT method with either the hard- or soft-prior scheme holds promise for in vivo acquiring the optical background of tissues, the low robustness of the hard-prior scheme to the segmentation error and inferior performance of the soft-prior one in the quantitative accuracy limit its further application. We propose in this paper a shape-parameterized DOT method for not only effectively determining the regional optical properties but potentially achieving reasonable structural amelioration, lending itself to FMT for comparably improved recovery of fluorescence distribution. PMID:25360379
Wang, Daifa; He, Jin; Qiao, Huiting; Song, Xiaolei; Fan, Yubo; Li, Deyu
2014-01-01
Fluorescence molecular tomography in the near-infrared region is becoming a powerful modality for mapping the three-dimensional quantitative distributions of fluorochromes in live small animals. However, wider application of fluorescence molecular tomography still requires more accurate and stable reconstruction tools. We propose a shape-based reconstruction method that uses spherical harmonics parameterization, where fluorophores are assumed to be distributed as piecewise constants inside disjointed subdomains and the remaining background. The inverse problem is then formulated as a constrained nonlinear least-squares problem with respect to shape parameters, which decreases ill-posedness because of the significantly reduced number of unknowns. Since different shape parameters contribute differently to the boundary measurements, a two-step and modified block coordinate descent optimization algorithm is introduced to stabilize the reconstruction. We first evaluated our method using numerical simulations under various conditions for the noise level and fluorescent background; it showed significant superiority over conventional voxel-based methods in terms of the spatial resolution, reconstruction accuracy with regard to the morphology and intensity, and robustness against the initial estimated distribution. In our phantom experiment, our method again showed better spatial resolution and more accurate intensity reconstruction. Finally, the results of an in vivo experiment demonstrated its applicability to the imaging of mice. PMID:24732826
NASA Technical Reports Server (NTRS)
Wentz, W. H., Jr.; Ostowari, C.
1983-01-01
Experimental measurements were made to determine the effects of slot gap opening and flap cove shape on flap and airfoil flow fields. Test model was the GA(W)-1 airfoil with 0.30c Fowler flap deflected 35 degrees. Tests were conducted with optimum, wide and narrow gaps, and with three cove shapes. Three test angles were selected, corresponding to pre-stall and post-stall conditions. Reynolds number was 2,200,000 and Mach number was 0.13. Force, surface pressure, total pressure, and split-film turbulence measurements were made. Results were compared with theory for those parameters for which theoretical values were available.
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.
Design of advanced airfoil for stall-regulated wind turbines
NASA Astrophysics Data System (ADS)
Grasso, F.; Coiro, D. P.; Bizzarrini, N.; Calise, G.
2016-09-01
Nowadays, all the modern MW-class wind turbines make use of pitch control to optimize the rotor performance and control the turbine. However, for kW-range machines, stall-regulated solutions are still attractive and largely used for their simplicity and robustness. On the design phase, the aerodynamics plays a crucial role, especially concerning the selection/design of the necessary airfoils. This is because the airfoil performance should guarantee high wind turbine performance, but also the needed machine control capabilities. In the present work, the design of a new airfoil dedicated for stall machines is discussed. The design strategy makes use of numerical optimization scheme where a gradient-based algorithm is coupled with XFOIL code and an original Bezier-curves-based parameterization to describe the airfoil shape. The performances of the new airfoil are compared in free and fixed transition conditions. In addition, the performance of the rotor is analysed comparing the impact of the new geometry with alternative candidates. The results show that the new airfoil offers better performance and control than existing candidates do.
NASA Technical Reports Server (NTRS)
Huyse, Luc; Bushnell, Dennis M. (Technical Monitor)
2001-01-01
Free-form shape optimization of airfoils poses unexpected difficulties. Practical experience has indicated that a deterministic optimization for discrete operating conditions can result in dramatically inferior performance when the actual operating conditions are different from the - somewhat arbitrary - design values used for the optimization. Extensions to multi-point optimization have proven unable to adequately remedy this problem of "localized optimization" near the sampled operating conditions. This paper presents an intrinsically statistical approach and demonstrates how the shortcomings of multi-point optimization with respect to "localized optimization" can be overcome. The practical examples also reveal how the relative likelihood of each of the operating conditions is automatically taken into consideration during the optimization process. This is a key advantage over the use of multipoint methods.
NASA Astrophysics Data System (ADS)
Thomareis, Nikitas; Papadakis, George
2017-01-01
Direct numerical simulations of the flow field around a NACA 0012 airfoil at Reynolds number 50 000 and angle of attack 5° with 3 different trailing edge shapes (straight, blunt, and serrated) have been performed. Both time-averaged flow characteristics and the most dominant flow structures and their frequencies are investigated using the dynamic mode decomposition method. It is shown that for the straight trailing edge airfoil, this method can capture the fundamental as well as the subharmonic of the Kelvin-Helmholtz instability that develops naturally in the separating shear layer. The fundamental frequency matches well with relevant data in the literature. The blunt trailing edge results in periodic vortex shedding, with frequency close to the subharmonic of the natural shear layer frequency. The shedding, resulting from a global instability, has an upstream effect and forces the separating shear layer. Due to forcing, the shear layer frequency locks onto the shedding frequency while the natural frequency (and its subharmonic) is suppressed. The presence of serrations in the trailing edge creates a spanwise pressure gradient, which is responsible for the development of a secondary flow pattern in the spanwise direction. This pattern affects the mean flow in the near wake. It can explain an unexpected observation, namely, that the velocity deficit downstream of a trough is smaller than the deficit after a protrusion. Furthermore, the insertion of serrations attenuates the energy of vortex shedding by de-correlating the spanwise coherence of the vortices. This results in weaker forcing of the separating shear layer, and both the subharmonics of the natural frequency and the shedding frequency appear in the spectra.
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
Large and Small Droplet Impingement Data on Airfoils and Two Simulated Ice Shapes
NASA Technical Reports Server (NTRS)
Papadakis, Michael; Wong, See-Cheuk; Rachman, Arief; Hung, Kuohsing E.; Vu, Giao T.; Bidwell, Colin S.
2007-01-01
Water droplet impingement data were obtained at the NASA Glenn Icing Research Tunnel (IRT) for four wings and one wing with two simulated ice shapes. The wings tested include three 36-in. chord wings (MS(1)-317, GLC-305, and a NACA 652-415) and a 57-in. chord Twin Otter horizontal tail section. The simulated ice shapes were 22.5- and 45-min glaze ice shapes for the Twin Otter horizontal tail section generated using the LEWICE 2.2 ice accretion program. The impingement experiments were performed with spray clouds having median volumetric diameters of 11, 21, 79, 137, and 168 mm. Comparisons to the experimental data were generated which showed good agreement for the clean wings and ice shapes at lower drop sizes. For larger drop sizes LEWICE 2.2 over predicted the collection efficiencies due to droplet splashing effects which were not modeled in the program. Also for the more complex glaze ice shapes interpolation errors resulted in the over prediction of collection efficiencies in cove and shadow regions of ice shapes.
Airfoil-shaped extension-twist-coupled composite star-beams for rotor blade tip applications
NASA Astrophysics Data System (ADS)
Mahadev, Sthanu
airfoil. An ABAQUS based finite element approach is employed to obtain a first characterization of the integral blade tip pitch response to changes in axial load and torque.
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.
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.
Tangler, J.L.; Somers, D.M.
2000-05-30
Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.
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.
Multiple piece turbine airfoil
Kimmel, Keith D
2010-11-09
A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of hook shaped struts each mounted within channels extending in a spanwise direction of the spar and the shell to allow for relative motion between the spar and shell in the airfoil chordwise direction while also fanning a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure. The hook struts have a hooked shaped end and a rounded shaped end in order to insert the struts into the spar.
Study of a new airfoil used in reversible axial fans
NASA Technical Reports Server (NTRS)
Li, Chaojun; Wei, Baosuo; Gu, Chuangang
1991-01-01
The characteristics of the reverse ventilation of axial flow are analyzed. An s shaped airfoil with a double circular arc was tested in a wind tunnel. The experimental results showed that the characteristics of this new airfoil in reverse ventilation are the same as those in normal ventilation, and that this airfoil is better than the existing airfoils used on reversible axial fans.
Morphing-Based Shape Optimization in Computational Fluid Dynamics
NASA Astrophysics Data System (ADS)
Rousseau, Yannick; Men'Shov, Igor; Nakamura, Yoshiaki
In this paper, a Morphing-based Shape Optimization (MbSO) technique is presented for solving Optimum-Shape Design (OSD) problems in Computational Fluid Dynamics (CFD). The proposed method couples Free-Form Deformation (FFD) and Evolutionary Computation, and, as its name suggests, relies on the morphing of shape and computational domain, rather than direct shape parameterization. Advantages of the FFD approach compared to traditional parameterization are first discussed. Then, examples of shape and grid deformations by FFD are presented. Finally, the MbSO approach is illustrated and applied through an example: the design of an airfoil for a future Mars exploration airplane.
Derbidge, Renatus; Feiten, Linus; Conradt, Oliver; Heusser, Peter; Baumgartner, Stephan
2013-01-01
Photographs of mistletoe (Viscum album L.) berries taken by a permanently fixed camera during their development in autumn were subjected to an outline shape analysis by fitting path curves using a mathematical algorithm from projective geometry. During growth and maturation processes the shape of mistletoe berries can be described by a set of such path curves, making it possible to extract changes of shape using one parameter called Lambda. Lambda describes the outline shape of a path curve. Here we present methods and software to capture and measure these changes of form over time. The present paper describes the software used to automatize a number of tasks including contour recognition, optimization of fitting the contour via hill-climbing, derivation of the path curves, computation of Lambda and blinding the pictures for the operator. The validity of the program is demonstrated by results from three independent measurements showing circadian rhythm in mistletoe berries. The program is available as open source and will be applied in a project to analyze the chronobiology of shape in mistletoe berries and the buds of their host trees.
Derbidge, Renatus; Feiten, Linus; Conradt, Oliver; Heusser, Peter; Baumgartner, Stephan
2013-01-01
Photographs of mistletoe (Viscum album L.) berries taken by a permanently fixed camera during their development in autumn were subjected to an outline shape analysis by fitting path curves using a mathematical algorithm from projective geometry. During growth and maturation processes the shape of mistletoe berries can be described by a set of such path curves, making it possible to extract changes of shape using one parameter called Lambda. Lambda describes the outline shape of a path curve. Here we present methods and software to capture and measure these changes of form over time. The present paper describes the software used to automatize a number of tasks including contour recognition, optimization of fitting the contour via hill-climbing, derivation of the path curves, computation of Lambda and blinding the pictures for the operator. The validity of the program is demonstrated by results from three independent measurements showing circadian rhythm in mistletoe berries. The program is available as open source and will be applied in a project to analyze the chronobiology of shape in mistletoe berries and the buds of their host trees. PMID:23565255
NASA Technical Reports Server (NTRS)
Spring, Samuel A.
1988-01-01
The flow field about a NACA 0012 airfoil with a simulated glaze ice shape was studied. Split hot-film anemometry was used to measure the streamwise velocity component in the upper and lower separation bubbles aft of the glaze ice horns. Velocity profiles were presented as well as the boundary layer momentum and displacement thickness distributions through the bubbles. Data were presented at angles of attack of 0, 2 and 4 deg and clearly showed the large region of reverse flow. A detailed discussion of split hot-film acquisition and data reduction, and possible sources of error has been included.
NASA Astrophysics Data System (ADS)
Bragg, M. B.; Broeren, A. P.; Blumenthal, L. A.
2005-07-01
Past research on airfoil aerodynamics in icing are reviewed. This review emphasizes the time period after the 1978 NASA Lewis workshop that initiated the modern icing research program at NASA and the current period after the 1994 ATR accident where aerodynamics research has been more aircraft safety focused. Research pre-1978 is also briefly reviewed. Following this review, our current knowledge of iced airfoil aerodynamics is presented from a flowfield-physics perspective. This article identifies four classes of ice accretions: roughness, horn ice, streamwise ice, and spanwise-ridge ice. For each class, the key flowfield features such as flowfield separation and reattachment are discussed and how these contribute to the known aerodynamic effects of these ice shapes. Finally Reynolds number and Mach number effects on iced-airfoil aerodynamics are summarized.
Flatback airfoil wind tunnel experiment.
Mayda, Edward A.; van Dam, C.P.; Chao, David D.; Berg, Dale E.
2008-04-01
A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.
Transonic single-mode flutter and buffet of a low aspect ratio wing having a subsonic airfoil shape
NASA Technical Reports Server (NTRS)
Erickson, L. L.
1974-01-01
Transonic flutter and buffet results obtained from wind-tunnel tests of a low aspect ratio semispan wing model are presented. The tests were conducted to investigate potential transonic aeroelastic problems of vehicles having subsonic airfoil sections. The model employed NACA 00XX-64 airfoil sections in the streamwise direction and had a 14 deg leading edge sweep angle. Aspect ratio, and average thickness were 4.0, 0.35, and 8 percent, respectively. The model was tested at Mach numbers from 0.6 to 0.95 at angles of attack from 0 deg to 15 deg. Two zero lift flutter conditions were found that involved essentially single normal mode vibrations. With boundary layer trips on the model, flutter occurred in a narrow Mach number range centered at about Mach 0.90. The frequency and motion of this flutter were like that of the first normal mode vibration. With the trips removed flutter occurred at a slightly high Mach number but in a mode strongly resembling that of the second normal mode.
NASA Technical Reports Server (NTRS)
Johnston, W. A.
1983-01-01
A viscous-inviscid interactive calculation procedure is developed for application to flow in cascades of two-dimensional airfoils. This procedure has essentially three components. First, a numerical solution of the Eulers equations which can accommodate an arbitrarily specified cascade geometry of the cascade. A method of grid generation has been used which relics in part on a succession of conformal mappings. Second, a viscous solution for use in boundary layers and wake regions was programmed. Finally, an interactive scheme which takes the form of a source-sink distribution along the blade surface and wake centerline is employed. Results were obtained with this procedure for several cascade flow situations, and some comparisons with experiment are presented.
Generalized multi-point inverse airfoil design
NASA Technical Reports Server (NTRS)
Selig, Michael S.; Maughmer, Mark D.
1991-01-01
In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution or boundary-layer development, etc., then from this information determine the corresponding airfoil shape. This paper presents a method which approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed. In addition to these local desired distributions, single parameters like the airfoil thickness can be specified. The problem of finding the airfoil shape is determined by coupling an incompressible, inviscid, inverse airfoil design method with a direct integral boundary-layer analysis method and solving the resulting nonlinear equations via a multidimensional Newton iteration technique. The approach is fast and easily allows for interactive design. It is also flexible and could be adapted to solving compressible, inverse airfoil design problems.
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.
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.
Garcia-Crespo, Andres Jose
2015-03-03
A composite blade assembly for mounting on a turbine wheel includes a ceramic airfoil and an airfoil platform. The ceramic airfoil is formed with an airfoil portion, a blade shank portion and a blade dovetail tang. The metal platform includes a platform shank and a radially inner platform dovetail. The ceramic airfoil is captured within the metal platform, such that in use, the ceramic airfoil is held within the turbine wheel independent of the metal platform.
Ice Accretions on Modern Airfoils Investigated
NASA Technical Reports Server (NTRS)
Addy, Harold E., Jr.
2000-01-01
The Icing Branch at the NASA Glenn Research Center at Lewis Field initiated and conducted the Modern Airfoils Ice Accretions project to identify ice shapes and determine their effects on the aerodynamic performance of aircraft, particularly on lift and drag. Previous aircraft ice shape and performance documentation focused on a few, older airfoils. This permitted more basic studies of the ice accretion process to be undertaken. However, having established both a working data base of ice shapes and the capability to predict these shapes for basic airfoils, questions arose about how ice might accrete differently on airfoils more representative of those being designed and flown on various aircraft today. Similarly, information about how these ice shapes would affect aerodynamic performance was needed.
NASA Technical Reports Server (NTRS)
Somers, Dan M. (Inventor)
2005-01-01
An airfoil having a fore airfoil element, an aft airfoil element, and a slot region in between them. These elements induce laminar flow over substantially all of the fore airfoil element and also provide for laminar flow in at least a portion of the slot region. The method of the invention is one for inducing natural laminar flow over an airfoil. In the method, a fore airfoil element, having a leading and trailing edge, and an aft airfoil element define a slot region. Natural laminar flow is induced over substantially all of the fore airfoil element, by inducing the pressures on both surfaces of the fore airfoil element to decrease to a location proximate the trailing edge of the fore airfoil element using pressures created by the aft airfoil element.
Supercritical Flow Past Symmetrical Airfoils.
1980-12-01
about quasi-elliptic airfoil sections. The method was later extended by Boerstoel [1967] to present a catalog of solutions for certain body shapes. Bauer...Lecture Notes in Economics and Mathematical Systems, Springer- Verlag, New York, 1972. Boerstoel , J. W., "A Survey of Symmetrical Transonic Potential
Modeling and Grid Generation of Iced Airfoils
NASA Technical Reports Server (NTRS)
Vickerman, Mary B.; Baez, Marivell; Braun, Donald C.; Hackenberg, Anthony W.; Pennline, James A.; Schilling, Herbert W.
2007-01-01
SmaggIce Version 2.0 is a software toolkit for geometric modeling and grid generation for two-dimensional, singleand multi-element, clean and iced airfoils. A previous version of SmaggIce was described in Preparing and Analyzing Iced Airfoils, NASA Tech Briefs, Vol. 28, No. 8 (August 2004), page 32. To recapitulate: Ice shapes make it difficult to generate quality grids around airfoils, yet these grids are essential for predicting ice-induced complex flow. This software efficiently creates high-quality structured grids with tools that are uniquely tailored for various ice shapes. SmaggIce Version 2.0 significantly enhances the previous version primarily by adding the capability to generate grids for multi-element airfoils. This version of the software is an important step in streamlining the aeronautical analysis of ice airfoils using computational fluid dynamics (CFD) tools. The user may prepare the ice shape, define the flow domain, decompose it into blocks, generate grids, modify/divide/merge blocks, and control grid density and smoothness. All these steps may be performed efficiently even for the difficult glaze and rime ice shapes. Providing the means to generate highly controlled grids near rough ice, the software includes the creation of a wrap-around block (called the "viscous sublayer block"), which is a thin, C-type block around the wake line and iced airfoil. For multi-element airfoils, the software makes use of grids that wrap around and fill in the areas between the viscous sub-layer blocks for all elements that make up the airfoil. A scripting feature records the history of interactive steps, which can be edited and replayed later to produce other grids. Using this version of SmaggIce, ice shape handling and grid generation can become a practical engineering process, rather than a laborious research effort.
Numerical Airfoil Optimization Using a Reduced Number of Design Coordinates
NASA Technical Reports Server (NTRS)
Vanderplaats, G. N.; Hicks, R. M.
1976-01-01
A method is presented for numerical airfoil optimization whereby a reduced number of design coordinates are used to define the airfoil shape. The approach is to define the airfoil as a linear combination of shapes. These basic shapes may be analytically or numerically defined, allowing the designer to use his insight to propose candidate designs. The design problem becomes one of determining the participation of each such function in defining the optimum airfoil. Examples are presented for two-dimensional airfoil design and are compared with previous results based on a polynomial representation of the airfoil shape. Four existing NACA airfoils are used as basic shapes. Solutions equivalent to previous results are achieved with a factor of more than 3 improvements in efficiency, while superior designs are demonstrated with an efficiency greater than 2 over previous methods. With this shape definition, the optimization process is shown to exploit the simplifying assumptions in the inviscid aerodynamic analysis used here, thus demonstrating the need to use more advanced aerodynamics for airfoil optimization.
Unsteady Airloads on Airfoils in Reverse Flow
NASA Astrophysics Data System (ADS)
Lind, Andrew; Jones, Anya
2014-11-01
This work gives insight into the influence of airfoil characteristics on unsteady airloads for rotor applications where local airfoil sections may operate at high and/or reverse flow angles of attack. Two-dimensional wind tunnel experiments have been performed on four airfoil sections to investigate the effects of thickness, camber, and trailing edge shape on unsteady airloads (lift, pressure drag, and pitching moment). These model rotor blades were tested through 360 deg of incidence for 104 <=Re <=106 . Unsteady pressure transducers were mounted on the airfoil surface to measure the high frequency, dynamic pressure variations. The temporal evolution of chordwise pressure distributions and resulting airloads is quantified for each airfoil in each of the three unsteady wake regimes present in reverse flow. Specifically, the influence of the formation, growth, and shedding of vortices on the surface pressure distribution is quantified and compared between airfoils with a sharp geometric trailing edge and those with a blunt geometric trailing edge. These findings are integral to mitigation of rotor blade vibrations for applications where airfoil sections are subjected to reverse flow, such as high-speed helicopters and tidal turbines.
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
1980-02-01
the elliptic cross section is considered to be more representative of the NACA 64A010 airfoil with boundary layer displacement thickness added on than...section and the flat plate airfoil with Kutta condition. The experimental results are for the NACA 64A010 airfoil at M = 0.5 and Reynolds number between...practice for actual airfoils. The experimental data shown in Fig. 3.5 are for the NACA 4 and 5 digit series airfoils (Ref. 17). The lift curve slope is
Simplified dragonfly airfoil aerodynamics at Reynolds numbers below 8000
NASA Astrophysics Data System (ADS)
Levy, David-Elie; Seifert, Avraham
2009-07-01
Effective aerodynamics at Reynolds numbers lower than 10 000 is of great technological interest and a fundamental scientific challenge. The current study covers a Reynolds number range of 2000-8000. At these Reynolds numbers, natural insect flight could provide inspiration for technology development. Insect wings are commonly characterized by corrugated airfoils. In particular, the airfoil of the dragonfly, which is able to glide, can be used for two-dimensional aerodynamic study of fixed rigid wings. In this study, a simplified dragonfly airfoil is numerically analyzed in a steady free-stream flow. The aerodynamic performance (such as mean and fluctuating lift and drag), are first compared to a "traditional" low Reynolds number airfoil: the Eppler-E61. The numerical results demonstrate superior performances of the corrugated airfoil. A series of low-speed wind and water tunnel experiments were performed on the corrugated airfoil, to validate the numerical results. The findings indicate quantitative agreement with the mean wake velocity profiles and shedding frequencies while validating the two dimensionality of the flow. A flow physics numerical study was performed in order to understand the underlying mechanism of corrugated airfoils at these Reynolds numbers. Airfoil shapes based on the flow field characteristics of the corrugated airfoil were built and analyzed. Their performances were compared to those of the corrugated airfoil, stressing the advantages of the latter. It was found that the flow which separates from the corrugations and forms spanwise vortices intermittently reattaches to the aft-upper arc region of the airfoil. This mechanism is responsible for the relatively low intensity of the vortices in the airfoil wake, reducing the drag and increasing the flight performances of this kind of corrugated airfoil as compared to traditional low Reynolds number airfoils such as the Eppler E-61.
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.
An analytical study for the design of advanced rotor airfoils
NASA Technical Reports Server (NTRS)
Kemp, L. D.
1973-01-01
A theoretical study has been conducted to design and evaluate two airfoils for helicopter rotors. The best basic shape, designed with a transonic hodograph design method, was modified to meet subsonic criteria. One airfoil had an additional constraint for low pitching-moment at the transonic design point. Airfoil characteristics were predicted. Results of a comparative analysis of helicopter performance indicate that the new airfoils will produce reduced rotor power requirements compared to the NACA 0012. The hodograph design method, written in CDC Algol, is listed and described.
Application of two procedures for dual-point design of transonic airfoils
NASA Technical Reports Server (NTRS)
Mineck, Raymond E.; Campbell, Richard L.; Allison, Dennis O.
1994-01-01
Two dual-point design procedures were developed to reduce the objective function of a baseline airfoil at two design points. The first procedure to develop a redesigned airfoil used a weighted average of the shapes of two intermediate airfoils redesigned at each of the two design points. The second procedure used a weighted average of two pressure distributions obtained from an intermediate airfoil redesigned at each of the two design points. Each procedure was used to design a new airfoil with reduced wave drag at the cruise condition without increasing the wave drag or pitching moment at the climb condition. Two cycles of the airfoil shape-averaging procedure successfully designed a new airfoil that reduced the objective function and satisfied the constraints. One cycle of the target (desired) pressure-averaging procedure was used to design two new airfoils that reduced the objective function and came close to satisfying the constraints.
Wind tunnel results of the high-speed NLF(1)-0213 airfoil
NASA Technical Reports Server (NTRS)
Sewall, William G.; Mcghee, Robert J.; Hahne, David E.; Jordan, Frank L., Jr.
1987-01-01
Wind tunnel tests were conducted to evaluate a natural laminar flow airfoil designed for the high speed jet aircraft in general aviation. The airfoil, designated as the High Speed Natural Laminar Flow (HSNLF)(1)-0213, was tested in two dimensional wind tunnels to investigate the performance of the basic airfoil shape. A three dimensional wing designed with this airfoil and a high lift flap system is also being evaluated with a full size, half span model.
A computer program for the design and analysis of low-speed airfoils
NASA Technical Reports Server (NTRS)
Eppler, R.; Somers, D. M.
1980-01-01
A conformal mapping method for the design of airfoils with prescribed velocity distribution characteristics, a panel method for the analysis of the potential flow about given airfoils, and a boundary layer method have been combined. With this combined method, airfoils with prescribed boundary layer characteristics can be designed and airfoils with prescribed shapes can be analyzed. All three methods are described briefly. The program and its input options are described. A complete listing is given as an appendix.
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)
Grid Sensitivity and Aerodynamic Optimization of Generic Airfoils
NASA Technical Reports Server (NTRS)
Sadrehaghighi, Ideen; Smith, Robert E.; Tiwari, Surendra N.
1995-01-01
An algorithm is developed to obtain the grid sensitivity with respect to design parameters for aerodynamic optimization. The procedure is advocating a novel (geometrical) parameterization using spline functions such as NURBS (Non-Uniform Rational B- Splines) for defining the airfoil geometry. An interactive algebraic grid generation technique is employed to generate C-type grids around airfoils. The grid sensitivity of the domain with respect to geometric design parameters has been obtained by direct differentiation of the grid equations. A hybrid approach is proposed for more geometrically complex configurations such as a wing or fuselage. The aerodynamic sensitivity coefficients are obtained by direct differentiation of the compressible two-dimensional thin-layer Navier-Stokes equations. An optimization package has been introduced into the algorithm in order to optimize the airfoil surface. Results demonstrate a substantially improved design due to maximized lift/drag ratio of the airfoil.
NASA Technical Reports Server (NTRS)
Ormsbee, A. I.
1977-01-01
Airfoil geometries were developed for low speed high lift applications, such as general aviation aircraft, propellers and helicopter rotors. The primary effort was to determine the extent to which the application of turbulent boundary layer separation criteria, plus manipulation of other input parameters, specifically trailing edging velocity ratio, could be utilized to achieve high C sub Lmax airfoils with relatively low drag at C sub Lmax. Both single-element and double-element airfoils were considered. Wind tunnel testing of some airfoils was included.
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.
Options for Robust Airfoil Optimization under Uncertainty
NASA Technical Reports Server (NTRS)
Padula, Sharon L.; Li, Wu
2002-01-01
A robust optimization method is developed to overcome point-optimization at the sampled design points. This method combines the best features from several preliminary methods proposed by the authors and their colleagues. The robust airfoil shape optimization is a direct method for drag reduction over a given range of operating conditions and has three advantages: (1) it prevents severe degradation in the off-design performance by using a smart descent direction in each optimization iteration, (2) it uses a large number of spline control points as design variables yet the resulting airfoil shape does not need to be smoothed, and (3) it allows the user to make a tradeoff between the level of optimization and the amount of computing time consumed. For illustration purposes, the robust optimization method is used to solve a lift-constrained drag minimization problem for a two-dimensional (2-D) airfoil in Euler flow with 20 geometric design variables.
Stiffness characteristics of airfoils under pulse loading
NASA Astrophysics Data System (ADS)
Turner, Kevin Eugene
The turbomachinery industry continually struggles with the adverse effects of contact rubs between airfoils and casings. The key parameter controlling the severity of a given rub event is the contact load produced when the airfoil tips incur into the casing. These highly non-linear and transient forces are difficult to calculate and their effects on the static and rotating components are not well understood. To help provide this insight, experimental and analytical capabilities have been established and exercised through an alliance between GE Aviation and The Ohio State University Gas Turbine Laboratory. One of the early findings of the program is the influence of blade flexibility on the physics of rub events. The core focus of the work presented in this dissertation is to quantify the influence of airfoil flexibility through a novel modeling approach that is based on the relationship between applied force duration and maximum tip deflection. This relationship is initially established using a series of forward, non-linear and transient analyses in which simulated impulse rub loads are applied. This procedure, although effective, is highly inefficient and costly to conduct by requiring numerous explicit simulations. To alleviate this issue, a simplified model, named the pulse magnification model, is developed that only requires a modal analysis and a static analyses to fully describe how the airfoil stiffness changes with respect to load duration. Results from the pulse magnification model are compared to results from the full transient simulation method and to experimental results, providing sound verification for the use of the modeling approach. Furthermore, a unique and highly efficient method to model airfoil geometries was developed and is outlined in this dissertation. This method produces quality Finite Element airfoil definitions directly from a fully parameterized mathematical model. The effectiveness of this approach is demonstrated by comparing modal
Approximate method of designing a two-element airfoil
NASA Astrophysics Data System (ADS)
Abzalilov, D. F.; Mardanov, R. F.
2011-09-01
An approximate method is proposed for designing a two-element airfoil. The method is based on reducing an inverse boundary-value problem in a doubly connected domain to a problem in a singly connected domain located on a multisheet Riemann surface. The essence of the method is replacement of channels between the airfoil elements by channels of flow suction and blowing. The shape of these channels asymptotically tends to the annular shape of channels passing to infinity on the second sheet of the Riemann surface. The proposed method can be extended to designing multielement airfoils.
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.
Uncertainty Quantification for Airfoil Icing
NASA Astrophysics Data System (ADS)
DeGennaro, Anthony Matteo
Ensuring the safety of airplane flight in icing conditions is an important and active arena of research in the aerospace community. Notwithstanding the research, development, and legislation aimed at certifying airplanes for safe operation, an analysis of the effects of icing uncertainties on certification quantities of interest is generally lacking. The central objective of this thesis is to examine and analyze problems in airfoil ice accretion from the standpoint of uncertainty quantification. We focus on three distinct areas: user-informed, data-driven, and computational uncertainty quantification. In the user-informed approach to uncertainty quantification, we discuss important canonical icing classifications and show how these categories can be modeled using a few shape parameters. We then investigate the statistical effects of these parameters. In the data-driven approach, we build statistical models of airfoil ice shapes from databases of actual ice shapes, and quantify the effects of these parameters. Finally, in the computational approach, we investigate the effects of uncertainty in the physics of the ice accretion process, by perturbing the input to an in-house numerical ice accretion code that we develop in this thesis.
NASA Technical Reports Server (NTRS)
Kohl, F. J.
1982-01-01
The methodology to predict deposit evolution (deposition rate and subsequent flow of liquid deposits) as a function of fuel and air impurity content and relevant aerodynamic parameters for turbine airfoils is developed in this research. The spectrum of deposition conditions encountered in gas turbine operations includes the mechanisms of vapor deposition, small particle deposition with thermophoresis, and larger particle deposition with inertial effects. The focus is on using a simplified version of the comprehensive multicomponent vapor diffusion formalism to make deposition predictions for: (1) simple geometry collectors; and (2) gas turbine blade shapes, including both developing laminar and turbulent boundary layers. For the gas turbine blade the insights developed in previous programs are being combined with heat and mass transfer coefficient calculations using the STAN 5 boundary layer code to predict vapor deposition rates and corresponding liquid layer thicknesses on turbine blades. A computer program is being written which utilizes the local values of the calculated deposition rate and skin friction to calculate the increment in liquid condensate layer growth along a collector surface.
Multiple piece turbine airfoil
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.
Using High Resolution Design Spaces for Aerodynamic Shape Optimization Under Uncertainty
NASA Technical Reports Server (NTRS)
Li, Wu; Padula, Sharon
2004-01-01
This paper explains why high resolution design spaces encourage traditional airfoil optimization algorithms to generate noisy shape modifications, which lead to inaccurate linear predictions of aerodynamic coefficients and potential failure of descent methods. By using auxiliary drag constraints for a simultaneous drag reduction at all design points and the least shape distortion to achieve the targeted drag reduction, an improved algorithm generates relatively smooth optimal airfoils with no severe off-design performance degradation over a range of flight conditions, in high resolution design spaces parameterized by cubic B-spline functions. Simulation results using FUN2D in Euler flows are included to show the capability of the robust aerodynamic shape optimization method over a range of flight conditions.
Closed loop steam cooled airfoil
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.
Tests of related forward-camber airfoils in the variable-density wind tunnel
NASA Technical Reports Server (NTRS)
Jacobs, Eastman N; Pinkerton, Robert M; Greenberg, Harry
1937-01-01
A recent investigation of numerous related airfoils indicated that positions of camber forward of the usual location resulted in an increase of the maximum lift. As an extension of this investigation, a series of forward-camber airfoils has been developed, the members of which show airfoil characteristics superior to those of the airfoils previously investigated. The primary object of this report is to present fully corrected results for airfoils in the useful range of shapes. With the data thus made available, an airplane designer may intelligently choose the best possible airfoil-section shape for a given application and may predict to a reasonable degree the aerodynamic characteristics to be expected in flight from the section shape chosen.
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.
Wind tunnel testing of low-drag airfoils
NASA Technical Reports Server (NTRS)
Harvey, W. Donald; Mcghee, R. J.; Harris, C. D.
1986-01-01
Results are presented for the measured performance recently obtained on several airfoil concepts designed to achieve low drag by maintaining extensive regions of laminar flow without compromising high-lift performance. The wind tunnel results extend from subsonic to transonic speeds and include boundary-layer control through shaping and suction. The research was conducted in the NASA Langley 8-Ft Transonic Pressure Tunnel (TPT) and Low Turbulence Pressure Tunnel (LTPT) which have been developed for testing such low-drag airfoils. Emphasis is placed on identifying some of the major factors influencing the anticipated performance of low-drag airfoils.
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
Computational design and analysis of flatback airfoil wind tunnel experiment.
Mayda, Edward A.; van Dam, C.P.; Chao, David D.; Berg, Dale E.
2008-03-01
A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.
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.
An assessment of airfoil design by numerical optimization
NASA Technical Reports Server (NTRS)
Hicks, R. M.; Murman, E. M.; Vanderplaats, G. N.
1974-01-01
A practical procedure for optimum design of aerodynamic shapes is demonstrated. The proposed procedure uses an optimization program based on the method of feasible directions coupled with an analysis program that uses a relaxation solution of the inviscid, transonic, small-disturbance equations. Results are presented for low-drag, nonlifting transonic airfoils. Extension of the method to lifting airfoils, other speed regimes, and to three dimensions if feasible.
TAIR- TRANSONIC AIRFOIL ANALYSIS COMPUTER CODE
NASA Technical Reports Server (NTRS)
Dougherty, F. C.
1994-01-01
The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters
Aerodynamic Simulation of Ice Accretion on Airfoils
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Addy, Harold E., Jr.; Bragg, Michael B.; Busch, Greg T.; Montreuil, Emmanuel
2011-01-01
This report describes recent improvements in aerodynamic scaling and simulation of ice accretion on airfoils. Ice accretions were classified into four types on the basis of aerodynamic effects: roughness, horn, streamwise, and spanwise ridge. The NASA Icing Research Tunnel (IRT) was used to generate ice accretions within these four types using both subscale and full-scale models. Large-scale, pressurized windtunnel testing was performed using a 72-in.- (1.83-m-) chord, NACA 23012 airfoil model with high-fidelity, three-dimensional castings of the IRT ice accretions. Performance data were recorded over Reynolds numbers from 4.5 x 10(exp 6) to 15.9 x 10(exp 6) and Mach numbers from 0.10 to 0.28. Lower fidelity ice-accretion simulation methods were developed and tested on an 18-in.- (0.46-m-) chord NACA 23012 airfoil model in a small-scale wind tunnel at a lower Reynolds number. The aerodynamic accuracy of the lower fidelity, subscale ice simulations was validated against the full-scale results for a factor of 4 reduction in model scale and a factor of 8 reduction in Reynolds number. This research has defined the level of geometric fidelity required for artificial ice shapes to yield aerodynamic performance results to within a known level of uncertainty and has culminated in a proposed methodology for subscale iced-airfoil aerodynamic simulation.
NASA Technical Reports Server (NTRS)
Hylton, Larry D.
1986-01-01
Emphasis is placed on developing more accurate analytical models for predicting turbine airfoil external heat transfer rates. Performance goals of new engines require highly refined, accurate design tools to meet durability requirements. In order to obtain improvements in analytical capabilities, programs are required which focus on enhancing analytical techniques through verification of new models by comparison with relevant experimental data. The objectives of the current program are to develop an analytical approach, based on boundary layer theory, for predicting the effects of airfoil film cooling on downstream heat transfer rates and to verify the resulting analytical method by comparison of predictions with hot cascade data obtained under this program.
NASA Technical Reports Server (NTRS)
Garabedian, P. R.
1979-01-01
Computer codes for the design and analysis of transonic airfoils are considered. The design code relies on the method of complex characteristics in the hodograph plane to construct shockless airfoil. The analysis code uses artificial viscosity to calculate flows with weak shock waves at off-design conditions. Comparisons with experiments show that an excellent simulation of two dimensional wind tunnel tests is obtained. The codes have been widely adopted by the aircraft industry as a tool for the development of supercritical wing technology.
2010-08-01
airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 4.) This characteristic is related to the...edge with increasing (decreasing) lift coefficient. This feature results in a leading-edge shape that produces a suction peak at higher lift...should look like sketch 3. Sketch 3 1Director, Institute for Aerodynamics and Gas Dynamics, University of Stuttgart, Germany, 1974–1985.5 No suction
Transonic airfoil and axial flow rotary machine
Nagai, Naonori; Iwatani, Junji
2015-09-01
Sectional profiles close to a tip 124 and a part between a midportion 125 and a hub 123 are shifted to the upstream of an operating fluid flow in a sweep direction. Accordingly, an S shape is formed in which the tip 124 and the part between the midportion 125 and the hub 123 protrude. As a result, it is possible reduce various losses due to shook, waves, thereby forming a transonic airfoil having an excellent aerodynamic characteristic.
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 Design and Rotorcraft Performance
NASA Technical Reports Server (NTRS)
Bousman, William G.
2003-01-01
The relationship between global performance of a typical helicopter and the airfoil environment, as represented by the airfoil angles of attack and Mach number, has been examined using the comprehensive analysis CAMRAD II. A general correspondence is observed between global performance parameters, such as rotor L/D, and airfoil performance parameters, such as airfoil L/D, the drag bucket boundaries, and the divergence Mach number. Effects of design parameters such as blade twist and rotor speed variation have been examined and, in most cases, improvements observed in global performance are also observed in terms of airfoil performance. The relations observed between global Performance and the airfoil environment suggests that the emphasis in airfoil design should be for good L/D, while the maximum lift coefficient performance is less important.
Airfoil Ice-Accretion Aerodynamics Simulation
NASA Technical Reports Server (NTRS)
Bragg, Michael B.; Broeren, Andy P.; Addy, Harold E.; Potapczuk, Mark G.; Guffond, Didier; Montreuil, E.
2007-01-01
NASA Glenn Research Center, ONERA, and the University of Illinois are conducting a major research program whose goal is to improve our understanding of the aerodynamic scaling of ice accretions on airfoils. The program when it is completed will result in validated scaled simulation methods that produce the essential aerodynamic features of the full-scale iced-airfoil. This research will provide some of the first, high-fidelity, full-scale, iced-airfoil aerodynamic data. An initial study classified ice accretions based on their aerodynamics into four types: roughness, streamwise ice, horn ice, and spanwise-ridge ice. Subscale testing using a NACA 23012 airfoil was performed in the NASA IRT and University of Illinois wind tunnel to better understand the aerodynamics of these ice types and to test various levels of ice simulation fidelity. These studies are briefly reviewed here and have been presented in more detail in other papers. Based on these results, full-scale testing at the ONERA F1 tunnel using cast ice shapes obtained from molds taken in the IRT will provide full-scale iced airfoil data from full-scale ice accretions. Using these data as a baseline, the final step is to validate the simulation methods in scale in the Illinois wind tunnel. Computational ice accretion methods including LEWICE and ONICE have been used to guide the experiments and are briefly described and results shown. When full-scale and simulation aerodynamic results are available, these data will be used to further develop computational tools. Thus the purpose of the paper is to present an overview of the program and key results to date.
Inverse airfoil design procedure using a multigrid Navier-Stokes method
NASA Technical Reports Server (NTRS)
Malone, J. B.; Swanson, R. C.
1991-01-01
The Modified Garabedian McFadden (MGM) design procedure was incorporated into an existing 2-D multigrid Navier-Stokes airfoil analysis method. The resulting design method is an iterative procedure based on a residual correction algorithm and permits the automated design of airfoil sections with prescribed surface pressure distributions. The new design method, Multigrid Modified Garabedian McFadden (MG-MGM), is demonstrated for several different transonic pressure distributions obtained from both symmetric and cambered airfoil shapes. The airfoil profiles generated with the MG-MGM code are compared to the original configurations to assess the capabilities of the inverse design method.
Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils
NASA Technical Reports Server (NTRS)
Lindsey, Walter F; Landrum, Emma Jean
1958-01-01
Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.
Method for forming a liquid cooled airfoil for a gas turbine
Grondahl, Clayton M.; Willmott, Leo C.; Muth, Myron C.
1981-01-01
A method for forming a liquid cooled airfoil for a gas turbine is disclosed. A plurality of holes are formed at spaced locations in an oversized airfoil blank. A pre-formed composite liquid coolant tube is bonded into each of the holes. The composite tube includes an inner member formed of an anti-corrosive material and an outer member formed of a material exhibiting a high degree of thermal conductivity. After the coolant tubes have been bonded to the airfoil blank, the airfoil blank is machined to a desired shape, such that a portion of the outer member of each of the composite tubes is contiguous with the outer surface of the machined airfoil blank. Finally, an external skin is bonded to the exposed outer surface of both the machined airfoil blank and the composite tubes.
An investigation of the aerodynamic characteristics of a new general aviation airfoil in flight
NASA Technical Reports Server (NTRS)
Gregorek, G. M.; Hoffmann, M. J.; Weislogel, G. S.
1982-01-01
A low speed airfoil, the GA(W)-2, - a 13% thickness to chord ratio airfoil was evaluated. The wing of a Beech Sundowner was modified at by adding balsa ribs and covered with aluminum skin, to alter the existing airfoil shape to that of the GA(W)-2 airfoil. The aircraft was flown in a flight test program that gathered wing surface pressures and wake data from which the lift drag, and pitching moment of the airfoil could be determined. After the base line performance of the airfoil was measured, the drag due to surface irregularities such as steps, rivets and surface waviness was determined. The potential reduction of drag through the use of surface coatings such as KAPTON was also investigated.
Vertical axis wind turbine airfoil
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.
Recent progress in the analysis of iced airfoils and wings
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer; Chen, Hsun H.; Kaups, Kalle; Schimke, Sue
1992-01-01
Recent work on the analysis of iced airfoils and wings is described. Ice shapes for multielement airfoils and wings are computed using an extension of the LEWICE code that was developed for single airfoils. The aerodynamic properties of the iced wing are determined with an interactive scheme in which the solutions of the inviscid flow equations are obtained from a panel method and the solutions of the viscous flow equations are obtained from an inverse three-dimensional finite-difference boundary-layer method. A new interaction law is used to couple the inviscid and viscous flow solutions. The newly developed LEWICE multielement code is amplified to a high-lift configuration to calculate the ice shapes on the slat and on the main airfoil and on a four-element airfoil. The application of the LEWICE wing code to the calculation of ice shapes on a MS-317 swept wing shows good agreement with measurements. The interactive boundary-layer method is applied to a tapered iced wing in order to study the effect of icing on the aerodynamic properties of the wing at several angles of attack.
Rime ice accretion and its effect on airfoil performance. Ph.D. Thesis. Final Report
NASA Technical Reports Server (NTRS)
Bragg, M. B.
1982-01-01
A methodology was developed to predict the growth of rime ice, and the resulting aerodynamic penalty on unprotected, subcritical, airfoil surfaces. The system of equations governing the trajectory of a water droplet in the airfoil flowfield is developed and a numerical solution is obtained to predict the mass flux of super cooled water droplets freezing on impact. A rime ice shape is predicted. The effect of time on the ice growth is modeled by a time-stepping procedure where the flowfield and droplet mass flux are updated periodically through the ice accretion process. Two similarity parameters, the trajectory similarity parameter and accumulation parameter, are found to govern the accretion of rime ice. In addition, an analytical solution is presented for Langmuir's classical modified inertia parameter. The aerodynamic evaluation of the effect of the ice accretion on airfoil performance is determined using an existing airfoil analysis code with empirical corrections. The change in maximum lift coefficient is found from an analysis of the new iced airfoil shape. The drag correction needed due to the severe surface roughness is formulated from existing iced airfoil and rough airfoil data. A small scale wind tunnel test was conducted to determine the change in airfoil performance due to a simulated rime ice shape.
NREL airfoil families for HAWTs
Tangler, J.L.; Somers, D.M.
1995-12-31
The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time nine airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c{sub 1,max}) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulated turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.
NREL airfoil families for HAWTs
Tangler, J L; Somers, D M
1995-01-01
The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time seven airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c{sub l,max}) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulated turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.
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.
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.
Turbine airfoil to shround attachment
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.
Lift enhancement of an airfoil using a Gurney flap and vortex generators
NASA Technical Reports Server (NTRS)
Storms, Bruce L.; Jang, Cory S.
1993-01-01
The results of a low-speed wind tunnel test are presented for a single-element airfoil incorporating two lift-enhancing devices, namely a Gurney flap and vortex generators. The former consists of a small plate, on the order of one to two percent of the airfoil chord in height, located at the trailing edge perpendicular to the pressure side of the airfoil. The later consist of commercially-available, wishbone-shaped vortex generators. The test was conducted in the NASA Ames 7- by 10-foot Wind Tunnel with a full-span NACA 4412 airfoil. Measurements of surface pressure distributions and wake profiles were made to determine the lift, drag, and pitching-moment coefficients for the various airfoil configurations. The results indicate that the addition of a Gurney flap increased the maximum lift coefficient from 1.49 up to 1.96.
Investigation of the Kline-Fogleman airfoil section for rotor blade applications
NASA Technical Reports Server (NTRS)
Lumsdaine, E.; Johnson, W. S.; Fletcher, L. M.; Peach, J. E.
1974-01-01
Wind tunnel tests of a wedgeshaped airfoil with sharp leading edge and a spanwise step were conducted. The airfoil was tested with variations of the following parameters: (1) Reynolds number, (2) step location, (3) step shape, (4) apex angle, and (5) with the step on either the upper or lower surface. The results are compared with a flat plate and with wedge airfoils without a step having the same aspect ratio. Water table tests were conducted for flow visualization and it was determined that the flow separates from the upper surface at low angles of attack. The wind tunnel tests show that the lift/drag ratio of the airfoil is lower than for a flat plate and the pressure data show that the airfoil derives its lift in the same manner as a flat plate.
Airfoil nozzle and shroud assembly
Shaffer, J.E.; Norton, P.F.
1997-06-03
An airfoil and nozzle assembly are disclosed including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached. 5 figs.
Airfoil nozzle and shroud assembly
Shaffer, James E.; Norton, Paul F.
1997-01-01
An airfoil and nozzle assembly including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached.
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.
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.
Analysis of a theoretically optimized transonic airfoil
NASA Technical Reports Server (NTRS)
Lores, M. E.; Burdges, K. P.; Shrewsbury, G. D.
1978-01-01
Numerical optimization was used in conjunction with an inviscid, full potential equation, transonic flow analysis computer code to design an upper surface contour for a conventional airfoil to improve its supercritical performance. The modified airfoil was tested in a compressible flow wind tunnel. The modified airfoil's performance was evaluated by comparison with test data for the baseline airfoil and for an airfoil developed by optimization of leading edge of the baseline airfoil. While the leading edge modification performed as expected, the upper surface re-design did not produce all of the expected performance improvements. Theoretical solutions computed using a full potential, transonic airfoil code corrected for viscosity were compared to experimental data for the baseline airfoil and the upper surface modification. These correlations showed that the theory predicted the aerodynamics of the baseline airfoil fairly well, but failed to accurately compute drag characteristics for the upper surface modification.
Nozzle airfoil having movable nozzle ribs
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.
Boundary Layer Control on Airfoils.
ERIC Educational Resources Information Center
Gerhab, George; Eastlake, Charles
1991-01-01
A phenomena, boundary layer control (BLC), produced when visualizing the fluidlike flow of air is described. The use of BLC in modifying aerodynamic characteristics of airfoils, race cars, and boats is discussed. (KR)
Second Stage Turbine Bucket Airfoil.
Xu, Liming; Ahmadi, Majid; Humanchuk, David John; Moretto, Nicholas; Delehanty, Richard Edward
2003-05-06
The second-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinate values defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.
Numerical design of shockless airfoils
NASA Technical Reports Server (NTRS)
Garabedian, P. R.
1979-01-01
An attempt is made to indicate and briefly discuss only the most significant achievements of the research. The most successful contribution from the contract was the code for two dimensional analysis of airfoils in transonic flow.
Thin airfoil theory based on approximate solution of the transonic flow equation
NASA Technical Reports Server (NTRS)
Spreiter, John R; Alksne, Alberta Y
1958-01-01
A method is presented for the approximate solution of the nonlinear equations of transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.
Thin airfoil theory based on approximate solution of the transonic flow equation
NASA Technical Reports Server (NTRS)
Spreiter, John R; Alksne, Alberta Y
1957-01-01
A method is presented for the approximate solution of the nonlinear equations transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.
NASA Technical Reports Server (NTRS)
Pulliam, T. H.; Nemec, M.; Holst, T.; Zingg, D. W.; Kwak, Dochan (Technical Monitor)
2002-01-01
A comparison between an Evolutionary Algorithm (EA) and an Adjoint-Gradient (AG) Method applied to a two-dimensional Navier-Stokes code for airfoil design is presented. Both approaches use a common function evaluation code, the steady-state explicit part of the code,ARC2D. The parameterization of the design space is a common B-spline approach for an airfoil surface, which together with a common griding approach, restricts the AG and EA to the same design space. Results are presented for a class of viscous transonic airfoils in which the optimization tradeoff between drag minimization as one objective and lift maximization as another, produces the multi-objective design space. Comparisons are made for efficiency, accuracy and design consistency.
Control of flow around a NACA 0012 airfoil with a micro-riblet film
NASA Astrophysics Data System (ADS)
Lee, S.-J.; Jang, Y.-G.
2005-07-01
The flow structure of the wake behind a NACA 0012 airfoil covered with a V-shaped micro-riblet film (hereafter, MRF) has been investigated experimentally. The results were compared with the corresponding results from an identical airfoil covered with a smooth polydimethylsiloxane (PDMS) film of the same thickness. The drag force acting on each airfoil, as well as the spatial distributions of turbulence statistics in the near wake behind each airfoil, were measured for Reynolds numbers (calculated based on the chord length, C=75mm) ranging from Re=1.03×104 to 5.14×104. At Re=1.54×104 (U0=3 m/s), the drag force on the MRF-covered airfoil was about 6.6% lower than that on the smooth airfoil. In contrast, at the higher Reynolds number of Re=4.62×104 (U0=9 m/s), application of the MRF increased the drag force by about 9.8%. To determine the spatial distributions of turbulence intensity, including the mean velocity, turbulence intensity and turbulent kinetic energy, 500 instantaneous velocity fields of the wake behind each airfoil were measured using a 2-frame PIV technique and ensemble averaged. For the case of drag reduction (Re=1.54×104), the near wake behind the MRF-covered airfoil had a shorter vortex formation region and higher vertical velocity component compared with that behind the smooth airfoil. At the downstream end of vortex formation region, the Reynolds shear stress and turbulent kinetic energy for the MRF-covered airfoil were similar or slightly larger than for the smooth airfoil. Smoke-wire flow visualization showed that the presence of the MRF on the airfoil surface caused the smoke filaments to become thinner and to be separated by a smaller lateral spacing, indicating suppression of spanwise movement. For the drag-increasing case (Re=4.62×104), the presence of MRF grooves on the airfoil seemed to increase the vertical velocity component and decrease the height of the large-scale streamwise vortices, which interacted actively. This active
Dynamic Stall Characteristics of Drooped Leading Edge Airfoils
NASA Technical Reports Server (NTRS)
Sankar, Lakshmi N.; Sahin, Mehmet; Gopal, Naveen
2000-01-01
Helicopters in high-speed forward flight usually experience large regions of dynamic stall over the retreating side of the rotor disk. The rapid variations in the lift and pitching moments associated with the stall process can result in vibratory loads, and can cause fatigue and failure of pitch links. In some instances, the large time lag between the aerodynamic forces and the blade motion can trigger stall flutter. A number of techniques for the alleviation of dynamic stall have been proposed and studied by researchers. Passive and active control techniques have both been explored. Passive techniques include the use of high solidity rotors that reduce the lift coefficients of individual blades, leading edge slots and leading edge slats. Active control techniques include steady and unsteady blowing, and dynamically deformable leading edge (DDLE) airfoils. Considerable amount of experimental and numerical data has been collected on the effectiveness of these concepts. One concept that has not received as much attention is the drooped-leading edge airfoil idea. It has been observed in wind tunnel studies and flight tests that drooped leading edge airfoils can have a milder dynamic stall, with a significantly milder load hysteresis. Drooped leading edge airfoils may not, however, be suitable at other conditions, e.g. in hover, or in transonic flow. Work needs to be done on the analysis and design of drooped leading edge airfoils for efficient operation in a variety of flight regimes (hover, dynamic stall, and transonic flow). One concept that is worthy of investigation is the dynamically drooping airfoil, where the leading edge shape is changed roughly once-per-rev to mitigate the dynamic stall.
Three-Dimensional Piecewise-Continuous Class-Shape Transformation of Wings
NASA Technical Reports Server (NTRS)
Olson, Erik D.
2015-01-01
Class-Shape Transformation (CST) is a popular method for creating analytical representations of the surface coordinates of various components of aerospace vehicles. A wide variety of two- and three-dimensional shapes can be represented analytically using only a modest number of parameters, and the surface representation is smooth and continuous to as fine a degree as desired. This paper expands upon the original two-dimensional representation of airfoils to develop a generalized three-dimensional CST parametrization scheme that is suitable for a wider range of aircraft wings than previous formulations, including wings with significant non-planar shapes such as blended winglets and box wings. The method uses individual functions for the spanwise variation of airfoil shape, chord, thickness, twist, and reference axis coordinates to build up the complete wing shape. An alternative formulation parameterizes the slopes of the reference axis coordinates in order to relate the spanwise variation to the tangents of the sweep and dihedral angles. Also discussed are methods for fitting existing wing surface coordinates, including the use of piecewise equations to handle discontinuities, and mathematical formulations of geometric continuity constraints. A subsonic transport wing model is used as an example problem to illustrate the application of the methodology and to quantify the effects of piecewise representation and curvature constraints.
Airfoil Vibration Dampers program
NASA Technical Reports Server (NTRS)
Cook, Robert M.
1991-01-01
The Airfoil Vibration Damper program has consisted of an analysis phase and a testing phase. During the analysis phase, a state-of-the-art computer code was developed, which can be used to guide designers in the placement and sizing of friction dampers. The use of this computer code was demonstrated by performing representative analyses on turbine blades from the High Pressure Oxidizer Turbopump (HPOTP) and High Pressure Fuel Turbopump (HPFTP) of the Space Shuttle Main Engine (SSME). The testing phase of the program consisted of performing friction damping tests on two different cantilever beams. Data from these tests provided an empirical check on the accuracy of the computer code developed in the analysis phase. Results of the analysis and testing showed that the computer code can accurately predict the performance of friction dampers. In addition, a valuable set of friction damping data was generated, which can be used to aid in the design of friction dampers, as well as provide benchmark test cases for future code developers.
A supercritical airfoil experiment
NASA Technical Reports Server (NTRS)
Mateer, G. G.; Seegmiller, H. L.; Hand, L. A.; Szodruck, J.
1994-01-01
The purpose of this investigation is to provide a comprehensive data base for the validation of numerical simulations. The objective of the present paper is to provide a tabulation of the experimental data. The data were obtained in the two-dimensional, transonic flowfield surrounding a supercritical airfoil. A variety of flows were studied in which the boundary layer at the trailing edge of the model was either attached or separated. Unsteady flows were avoided by controlling the Mach number and angle of attack. Surface pressures were measured on both the model and wind tunnel walls, and the flowfield surrounding the model was documented using a laser Doppler velocimeter (LDV). Although wall interference could not be completely eliminated, its effect was minimized by employing the following techniques. Sidewall boundary layers were reduced by aspiration, and upper and lower walls were contoured to accommodate the flow around the model and the boundary-layer growth on the tunnel walls. A data base with minimal interference from a tunnel with solid walls provides an ideal basis for evaluating the development of codes for the transonic speed range because the codes can include the wall boundary conditions more precisely than interference connections can be made to the data sets.
Root region airfoil for wind turbine
Tangler, James L.; Somers, Dan M.
1995-01-01
A thick airfoil for the root region of the blade of a wind turbine. The airfoil has a thickness in a range from 24%-26% and a Reynolds number in a range from 1,000,000 to 1,800,000. The airfoil has a maximum lift coefficient of 1.4-1.6 that has minimum sensitivity to roughness effects.
Advanced technology airfoil research, volume 2. [conferences
NASA Technical Reports Server (NTRS)
1979-01-01
A comprehensive review of airfoil research is presented. The major thrust of the research is in three areas: development of computational aerodynamic codes for airfoil analysis and design, development of experimental facilities and test techniques, and all types of airfoil applications.
Parameterizing turbulence over abrupt topography
NASA Astrophysics Data System (ADS)
Klymak, Jody
2016-11-01
Stratified flow over abrupt topography generates a spectrum of propagating internal waves at large scales, and non-linear overturning breaking waves at small scales. For oscillating flows, the large scale waves propagate away as internal tides, for steady flows the large-scale waves propagate away as standing "columnar modes". At small-scales, the breaking waves appear to be similar for either oscillating or steady flows, so long as in the oscillating case the topography is significantly steeper than the internal tide angle of propagation. The size and energy lost to the breaking waves can be predicted relatively well from assuming that internal modes that propagate horizontally more slowly than the barotropic internal tide speed are arrested and their energy goes to turbulence. This leads to a recipe for dissipation of internal tides at abrupt topography that is quite robust for both the local internal tide generation problem (barotropic forcing) and for the scattering problem (internal tides incident on abrupt topography). Limitations arise when linear generation models break down, an example of which is interference between two ridges. A single "super-critical" ridge is well-modeled by a single knife-edge topography, regardless of its actual shape, but two supercritical ridges in close proximity demonstrate interference of the high modes that makes knife-edfe approximations invalid. Future direction of this research will be to use more complicated linear models to estimate the local dissipation. Of course, despite the large local dissipation, many ridges radiate most of their energy into the deep ocean, so tracking this low-mode radiated energy is very important, particularly as it means dissipation parameterizations in the open ocean due to these sinks from the surface tide cannot be parameterized locally to where they are lost from the surface tide, but instead lead to non-local parameterizations. US Office of Naval Research; Canadian National Science and
Application of numerical optimization to the design of advanced supercritical airfoils
NASA Technical Reports Server (NTRS)
Johnson, R. R.; Hicks, R. M.
1979-01-01
An application of numerical optimization to the design of advanced airfoils for transonic aircraft showed that low-drag sections can be developed for a given design Mach number without an accompanying drag increase at lower Mach numbers. This is achieved by imposing a constraint on the drag coefficient at an off-design Mach number while minimizing the drag coefficient at the design Mach number. This multiple design-point numerical optimization has been implemented with the use of airfoil shape functions which permit a wide range of attainable profiles during the optimization process. Analytical data for the starting airfoil shape, a single design-point optimized shape, and a double design-point optimized shape are presented. Experimental data obtained in the NASA Ames two-by two-foot wind tunnel are also presented and discussed.
Advanced Surface Flux Parameterization
2001-09-30
within PE 0602435N are BE-35-2-18, for the Mesoscale Modeling of the Atmos- phere and Aerosols, BE-35-2-19, and for the Exploratory Data Assimilation ... Methods . Related project at NPS is N0001401WR20242 for Evaluating Surface Flux and Boundary Layer Parameterizations in Mesoscale Models Using
Stochastic Convection Parameterizations
NASA Technical Reports Server (NTRS)
Teixeira, Joao; Reynolds, Carolyn; Suselj, Kay; Matheou, Georgios
2012-01-01
computational fluid dynamics, radiation, clouds, turbulence, convection, gravity waves, surface interaction, radiation interaction, cloud and aerosol microphysics, complexity (vegetation, biogeochemistry, radiation versus turbulence/convection stochastic approach, non-linearities, Monte Carlo, high resolutions, large-Eddy Simulations, cloud structure, plumes, saturation in tropics, forecasting, parameterizations, stochastic, radiation-clod interaction, hurricane forecasts
Dynamic Parameterization of IPSEC
2001-12-01
EXPECTED BENFITS OF THE RESEARCH ............................................2 D. RESEARCH OBJECTIVES...182 3. Explore Proposal Caching Issues ...................................................182 4. Security Policy Editor...C. EXPECTED BENFITS OF THE RESEARCH By providing dynamic parameterization to IPsec, government and military 3 security systems will be able to
Hook nozzle arrangement for supporting airfoil vanes
Shaffer, James E.; Norton, Paul F.
1996-01-01
A gas turbine engine's nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic.
Effects of enviromentally imposed roughness on airfoil performance
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer
1987-01-01
The experimental evidence for the effects of rain, insects, and ice on airfoil performance are examined. The extent to which the available information can be incorporated in a calculation method in terms of change of shape and surface roughness is discussed. The methods described are based on the interactive boundary layer procedure of Cebeci or on the thin layer Navier Stokes procedure developed at NASA. Cases presented show that extensive flow separation occurs on the rough surfaces.
Effects of environmentally imposed roughness on airfoil performance
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer
1987-01-01
The experimental evidence for the effects of rain, insects, and ice on airfoil performance are examined. The extent to which the available information can be incorporated in a calculation method in terms of change of shape and surface roughness is discussed. The methods described are based on the interactive boundary procedure of Cebeci or on the thin layer Navier Stokes procedure developed at NASA. Cases presented show that extensive flow separation occurs on the rough surfaces.
Lightweight Parameterized Suffix Array Construction
NASA Astrophysics Data System (ADS)
Tomohiro, I.; Deguchi, Satoshi; Bannai, Hideo; Inenaga, Shunsuke; Takeda, Masayuki
We present a first algorithm for direct construction of parameterized suffix arrays and parameterized longest common prefix arrays for non-binary strings. Experimental results show that our algorithm is much faster than naïve methods.
Pneumatic Spoiler Controls Airfoil Lift
NASA Technical Reports Server (NTRS)
Hunter, D.; Krauss, T.
1991-01-01
Air ejection from leading edge of airfoil used for controlled decrease of lift. Pneumatic-spoiler principle developed for equalizing lift on helicopter rotor blades. Also used to enhance aerodynamic control of short-fuselage or rudderless aircraft such as "flying-wing" airplanes. Leading-edge injection increases maneuverability of such high-performance fixed-wing aircraft as fighters.
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
Partwise cross-parameterization via nonregular convex hull domains.
Wu, Huai-Yu; Pan, Chunhong; Zha, Hongbin; Yang, Qing; Ma, Songde
2011-10-01
In this paper, we propose a novel partwise framework for cross-parameterization between 3D mesh models. Unlike most existing methods that use regular parameterization domains, our framework uses nonregular approximation domains to build the cross-parameterization. Once the nonregular approximation domains are constructed for 3D models, different (and complex) input shapes are transformed into similar (and simple) shapes, thus facilitating the cross-parameterization process. Specifically, a novel nonregular domain, the convex hull, is adopted to build shape correspondence. We first construct convex hulls for each part of the segmented model, and then adopt our convex-hull cross-parameterization method to generate compatible meshes. Our method exploits properties of the convex hull, e.g., good approximation ability and linear convex representation for interior vertices. After building an initial cross-parameterization via convex-hull domains, we use compatible remeshing algorithms to achieve an accurate approximation of the target geometry and to ensure a complete surface matching. Experimental results show that the compatible meshes constructed are well suited for shape blending and other geometric applications.
OUT Success Stories: Advanced Airfoils for Wind Turbines
DOE R&D Accomplishments Database
Jones, J.; Green, B.
2000-08-01
New airfoils have substantially increased the aerodynamic efficiency of wind turbines. It is clear that these new airfoils substantially increased energy output from wind turbines. Virtually all new blades built in this country today use these advanced airfoil designs.
Thin oblique airfoils at supersonic speed
NASA Technical Reports Server (NTRS)
Jone, Robert T
1946-01-01
The well-known methods of thin-airfoil theory have been extended to oblique or sweptback airfoils of finite aspect ratio moving at supersonic speeds. The cases considered thus far are symmetrical airfoils at zero lift having plan forms bounded by straight lines. Because of the conical form of the elementary flow fields, the results are comparable in simplicity to the results of the two-dimensional thin-airfoil theory for subsonic speeds. In the case of untapered airfoils swept back behind the Mach cone the pressure distribution at the center section is similar to that given by the Ackeret theory for a straight airfoil. With increasing distance from the center section the distribution approaches the form given by the subsonic-flow theory. The pressure drag is concentrated chiefly at the center section and for long wings a slight negative drag may appear on outboard sections. (author)
A Theory of Unstaggered Airfoil Cascades in Compressible Flow
NASA Technical Reports Server (NTRS)
Spurr, Robert A.; Allen, H. Julian
1947-01-01
By use of the methods of thin airfoil theory, which include effects of compressibility, rela.tio^as are developed which permit the rapid determination of the pressure distribution over an unstaggered cascade of airfoils of a given profile, and the determination of the profile shape necessary to yield a given pressure distribution for small chord gap ratios, For incompressible flow the results of the theory are compared with available examples obtained by the more exact method of conformal transformation. Although the theory is developed for small chord/gap ratios, these comparisons show that it may be extended to chord/gap ratios of order unity, at least for low speed flows. Choking of cascades, a phenomenon of particular importance in compressor design, is considered.
A theory of unstaggered airfoil cascades in compressible flow
NASA Technical Reports Server (NTRS)
Spurr, Robert A; Allen, H Julian
1947-01-01
By use of the methods of thin airfoil theory, which include effects of compressibility, relations are developed which permit the rapid determination of the pressure distribution over an unstaggered cascade of airfoils of a given profile, and the determination of the profile shape necessary to yield a given pressure distribution for small chord/gap ratios. For incompressible flow the results of the theory are compared with available examples obtained by the more exact method of conformal transformation. Although the theory is developed for small chord/gap ratios, these comparisons show that it may be extended to chord/gap ratios of order unity, at least for low-speed flows. Choking cascades, a phenomenon of particular importance in compressor design, is considered.
PROFILE: Airfoil Geometry Manipulation and Display. User's Guide
NASA Technical Reports Server (NTRS)
Collins, Leslie; Saunders, David
1997-01-01
This report provides user information for program PROFILE, an aerodynamics design utility for plotting, tabulating, and manipulating airfoil profiles. A dozen main functions are available. The theory and implementation details for two of the more complex options are also presented. These are the REFINE option, for smoothing curvature in selected regions while retaining or seeking some specified thickness ratio, and the OPTIMIZE option, which seeks a specified curvature distribution. Use of programs QPLOT and BPLOT is also described, since all of the plots provided by PROFILE (airfoil coordinates, curvature distributions, pressure distributions)) are achieved via the general-purpose QPLOT utility. BPLOT illustrates (again, via QPLOT) the shape functions used by two of PROFILE's options. These three utilities should be distributed as one package. They were designed and implemented for the Applied Aerodynamics Branch at NASA Ames Research Center, Moffett Field, California. They are all written in FORTRAN 77 and run on DEC and SGI systems under OpenVMS and IRIX.
Control theory based airfoil design using the Euler equations
NASA Technical Reports Server (NTRS)
Jameson, Antony; Reuther, James
1994-01-01
This paper describes the implementation of optimization techniques based on control theory for airfoil design. In our previous work it was shown that control theory could be employed to devise effective optimization procedures for two-dimensional profiles by using the potential flow equation with either a conformal mapping or a general coordinate system. The goal of our present work is to extend the development to treat the Euler equations in two-dimensions by procedures that can readily be generalized to treat complex shapes in three-dimensions. Therefore, we have developed methods which can address airfoil design through either an analytic mapping or an arbitrary grid perturbation method applied to a finite volume discretization of the Euler equations. Here the control law serves to provide computationally inexpensive gradient information to a standard numerical optimization method. Results are presented for both the inverse problem and drag minimization problem.
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
Root region airfoil for wind turbine
Tangler, J.L.; Somers, D.M.
1995-05-23
A thick airfoil is described for the root region of the blade of a wind turbine. The airfoil has a thickness in a range from 24%--26% and a Reynolds number in a range from 1,000,000 to 1,800,000. The airfoil has a maximum lift coefficient of 1.4--1.6 that has minimum sensitivity to roughness effects. 3 Figs.
Boundary-layer stability and airfoil design
NASA Technical Reports Server (NTRS)
Viken, Jeffrey K.
1986-01-01
Several different natural laminar flow (NLF) airfoils have been analyzed for stability of the laminar boundary layer using linear stability codes. The NLF airfoils analyzed come from three different design conditions: incompressible; compressible with no sweep; and compressible with sweep. Some of the design problems are discussed, concentrating on those problems associated with keeping the boundary layer laminar. Also, there is a discussion on how a linear stability analysis was effectively used to improve the design for some of the airfoils.
Airfoil seal system for gas turbine engine
None, None
2013-06-25
A turbine airfoil seal system of a turbine engine having a seal base with a plurality of seal strips extending therefrom for sealing gaps between rotational airfoils and adjacent stationary components. The seal strips may overlap each other and may be generally aligned with each other. The seal strips may flex during operation to further reduce the gap between the rotational airfoils and adjacent stationary components.
Evaluation of a stalled airfoil analysis program
NASA Technical Reports Server (NTRS)
Rumsey, C. L.
1985-01-01
The Stalled Airfoil Analysis Program (SAAP) is a computer code for predicting the aerodynamic characteristics of an airfoil up to, and beyond, stall. SAAP is presently evaluated through comparisons with experiments and with two other theoretical methods over an extensive range of airfoils and Reynolds number conditions. SAAP modeled drag more accurately than either of the other methods, and at angles of attack below stall yielded a smoother lift variation with angle of attack.
NASA Astrophysics Data System (ADS)
Mughal, Umair Najeeb
2017-01-01
Flow around an airfoil to calculate pressure co-efficient variations at different relative velocities have always been an important/basic part of Aerodynamic Study. Potential flow theory is used to study flow behavior on rankine half body, non-rotating cylinder and rotating cylinder as it is more trackable. Falkan-Skan Similarity Solution is taken to simulate the flow behavior on wedge. However, to use potential flow theory on usable airfoils the author have used conformal mapping to show a relation between realistic airfoil shapes and the knowledge gained from flow about cylinders. This method can further be used in the designing of an airfoil section. The author has used Joukowski Tranform to generate the flow around airfoils of various geometries and then utilized Kutta condition to force the stagnation point at the trailing edge. Co-efficient of pressure over the entire airfoil surface were calculated and corrected using Karman-Tsien compressibility correction equations. On the basis of this, the location of the ports to install the flush measurement system is suggested.
Robust Airfoil Optimization to Achieve Consistent Drag Reduction Over a Mach Range
NASA Technical Reports Server (NTRS)
Li, Wu; Huyse, Luc; Padula, Sharon; Bushnell, Dennis M. (Technical Monitor)
2001-01-01
We prove mathematically that in order to avoid point-optimization at the sampled design points for multipoint airfoil optimization, the number of design points must be greater than the number of free-design variables. To overcome point-optimization at the sampled design points, a robust airfoil optimization method (called the profile optimization method) is developed and analyzed. This optimization method aims at a consistent drag reduction over a given Mach range and has three advantages: (a) it prevents severe degradation in the off-design performance by using a smart descent direction in each optimization iteration, (b) there is no random airfoil shape distortion for any iterate it generates, and (c) it allows a designer to make a trade-off between a truly optimized airfoil and the amount of computing time consumed. For illustration purposes, we use the profile optimization method to solve a lift-constrained drag minimization problem for 2-D airfoil in Euler flow with 20 free-design variables. A comparison with other airfoil optimization methods is also included.
Design of a family of new advanced airfoils for low wind class turbines
NASA Astrophysics Data System (ADS)
Grasso, Francesco
2014-12-01
In order to maximize the ratio of energy capture and reduce the cost of energy, the selection of the airfoils to be used along the blade plays a crucial role. Despite the general usage of existing airfoils, more and more, families of airfoils specially tailored for specific applications are developed. The present research is focused on the design of a new family of airfoils to be used for the blade of one megawatt wind turbine working in low wind conditions. A hybrid optimization scheme has been implemented, combining together genetic and gradient based algorithms. Large part of the work is dedicated to present and discuss the requirements that needed to be satisfied in order to have a consistent family of geometries with high efficiency, high lift and good structural characteristics. For each airfoil, these characteristics are presented and compared to the ones of existing airfoils. Finally, the aerodynamic design of a new blade for low wind class turbine is illustrated and compared to a reference shape developed by using existing geometries. Due to higher lift performance, the results show a sensitive saving in chords, wetted area and so in loads in idling position.
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.
Inverse transonic airfoil design including viscous interaction
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1976-01-01
A numerical technique was developed for the analysis of specified transonic airfoils or for the design of airfoils having a prescribed pressure distribution, including the effect of weak viscous interaction. The method uses the full potential equation, a stretched Cartesian coordinate system, and the Nash-MacDonald turbulent boundary layer method. Comparisons with experimental data for typical transonic airfoils show excellent agreement. An example shows the application of the method to design a thick aft-cambered airfoil, and the effects of viscous interaction on its performance are discussed.
Transonic airfoil flowfield analysis using Cartesian coordinates
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1975-01-01
A numerical technique for analyzing transonic airfoils is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that Cartesian coordinates are used rather than a grid which fits the airfoil, such as the conformal circle-plane or 'sheared parabolic' coordinates which were used previously. Comparison with previous results shows that it is not necessary to match the computational grid to the airfoil surface, and that accurate results can be obtained with a Cartesian grid for lifting supercritical airfoils.
Wavy flow cooling concept for turbine airfoils
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.
Schooling behavior of heaving flexible airfoils
NASA Astrophysics Data System (ADS)
Im, Sunghyuk; Sung, Hyung Jin
2016-11-01
The schooling behavior of rigid and flexible NACA0017 airfoils in the heaving motion is experimentally explored in a merry-go-round equipment. The airfoil was attached to the end of a horizontal support bar whose other end was connected to the freely rotating vertical axis. The axis was forced to undergo a sinusoidal motion in the vertical direction to make a pure heaving motion of the airfoils in the frequency range of 0.5 to 5 Hz. The propulsion due to the heaving airfoils is expressed by a horizontally rotating speed of the support bar. This experimental setup is simulating infinite schooling situations of airfoils in an in-phase heaving motion with the streamwise distance d. The ratio of the distance to the chord length d/ c was determined by the number of airfoils (1 <= n <= 8) . The rotational frequency F according to the heaving frequency f was measured with different experimental parameters. The schooling number S = f /(nF), representing the number of heaving oscillations between each airfoil, was introduced to explain the schooling behavior of the airfoils. The effects of the flexibility, d/ c and f on the propulsive performance were examined with the schooling behavior of the airfoils. This work was supported by the Creative Research Initiatives (No. 2016-004749) program of the National Research Foundation of Korea (MSIP).
Plasma Flow Control Optimized Airfoil
NASA Astrophysics Data System (ADS)
Voikov, Vladimir; Patel, Mehul
2005-11-01
Recent advances in flow control research have demonstrated that plasma actuators can be efficient in different aerodynamic applications, particularly in providing flight control without conventional moving surfaces. The concept involves the use of a laminar airfoil design that employs a separation ramp at the trailing edge that can be manipulated by a plasma actuator to control lift, similar to trailing-edge flaps. The advantages are lower drag by a combination of the laminar flow design, and elimination of parasitic drag associated with wing-flap junctions. This work involves numerical simulations and experiments on a HSNLF(1)-0213 airfoil. The numerical results are obtained using an unsteady, compressible Navier-Stokes simulation that includes a model for the plasma actuators. The experiments are performed on a 2-D airfoil section that is mounted on a lift-drag force balance. The results demonstrate lift enhancement produced by the plasma actuator that is comparable to a plane flap. They also reveal an optimum actuator unsteady frequency that scales with the length of the separated region and local velocity, and is associated with the generation of a train of spanwise vortices. Other scaling including the effect of Reynolds number is presented.
Icing Test Results on an Advanced Two-Dimensional High-Lift Multi-Element Airfoil
NASA Technical Reports Server (NTRS)
Shin, Jaiwon; Wilcox, Peter; Chin, Vincent; Sheldon, David
1994-01-01
An experimental study has been conducted to investigate ice accretions on a high-lift, multi-element airfoil in the Icing Research Tunnel at the NASA Lewis Research Center. The airfoil is representative of an advanced transport wing design. The experimental work was conducted as part of a cooperative program between McDonnell Douglas Aerospace and the NASA Lewis Research Center to improve current understanding of ice accretion characteristics on the multi-element airfoil. The experimental effort also provided ice shapes for future aerodynamic tests at flight Reynolds numbers to ascertain high-lift performance effects. Ice shapes documented for a landing configuration over a variety of icing conditions are presented along with analyses.
NASA Technical Reports Server (NTRS)
Harris, C. D.
1971-01-01
Wind-tunnel tests have been conducted at Mach numbers from 0.60 to 0.81 to determine the effects of trailing-edge geometry on the aerodynamic characteristics of a NASA supercritical airfoil shape. Variations in trailing-edge thicknesses from 0 to 1.5 percent of the chord and a cavity in the trailing edge were investigated with airfoils with maximum thicknesses of 10 and 11 percent of the chord.
Two-dimensional cascade test of a highly loaded, low-solidity, tandem airfoil turbine rotor blade
NASA Technical Reports Server (NTRS)
Kline, J. F.; Stabe, R. G.
1973-01-01
A tip region section of a low-solidity tandem airfoil blade for a turbine rotor was tested in a two-dimensional cascade tunnel at solidities of 0.736 and 0.912. Blade surface static pressures and blade exit total and static pressure and flow angle were surveyed. Blade surface velocities, wake shapes, and kinetic energy losses were analyzed and compared with values for 1.852 solidity tandem airfoil blading.
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.
Prediction of Film Cooling on Gas Turbine Airfoils
NASA Technical Reports Server (NTRS)
Garg, Vijay K.; Gaugler, Raymond E.
1994-01-01
A three-dimensional Navier-Stokes analysis tool has been developed in order to study the effect of film cooling on the flow and heat transfer characteristics of actual turbine airfoils. An existing code (Arnone et al., 1991) has been modified for the purpose. The code is an explicit, multigrid, cell-centered, finite volume code with an algebraic turbulence model. Eigenvalue scaled artificial dissipation and variable-coefficient implicit residual smoothing are used with a full-multigrid technique. Moreover, Mayle's transition criterion (Mayle, 1991) is used. The effects of film cooling have been incorporated into the code in the form of appropriate boundary conditions at the hole locations on the airfoil surface. Each hole exit is represented by several control volumes, thus providing an ability to study the effect of hole shape on the film-cooling characteristics. Comparison is fair with near mid-span experimental data for four and nine rows of cooling holes, five on the shower head, and two rows each on the pressure and suction surfaces. The computations, however, show a strong spanwise variation of the heat transfer coefficient on the airfoil surface, specially with shower-head cooling.
Propulsion of a flapping and oscillating airfoil
NASA Technical Reports Server (NTRS)
Garrick, I E
1937-01-01
Formulas are given for the propelling or drag force experience in a uniform air stream by an airfoil or an airfoil-aileron combination, oscillating in any of three degrees of freedom; vertical flapping, torsional oscillations about a fixed axis parallel to the span, and angular oscillations of the aileron about a hinge.
AFSMO/AFSCL- AIRFOIL SMOOTHING AND SCALING
NASA Technical Reports Server (NTRS)
Morgan, H. L
1994-01-01
Since its early beginnings, NASA has been actively involved in the design and testing of airfoil sections for a wide variety of applications. Recently a set of programs has been developed to smooth and scale arbitrary airfoil coordinates. The smoothing program, AFSMO, utilizes both least-squares polynomial and least-squares cubic-spline techniques to iteratively smooth the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. The camber and thickness distribution of the smooth airfoil are also computed. The scaling program, AFSCL, may then be used to scale the thickness distribution generated by the smoothing program to a specified maximum thickness. Once the thickness distribution has been scaled, it is combined with the camber distribution to obtain the final scaled airfoil contour. The airfoil smoothing and scaling programs are written in FORTRAN IV for batch execution and have been implemented on a CDC CYBER 170 series computer with a central memory requirement of approximately 70K (octal) of 60 bit words. Both programs generate plotted output via CALCOMP type plotting calls. These programs were developed in 1983.
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.
Measuring Lift with the Wright Airfoils
ERIC Educational Resources Information Center
Heavers, Richard M.; Soleymanloo, Arianne
2011-01-01
In this laboratory or demonstration exercise, we mount a small airfoil with its long axis vertical at one end of a nearly frictionless rotating platform. Air from a leaf blower produces a sidewise lift force L on the airfoil and a drag force D in the direction of the air flow (Fig. 1). The rotating platform is kept in equilibrium by adding weights…
NASA Astrophysics Data System (ADS)
Manela, A.
2016-07-01
The acoustic signature of an acoustically compact tandem airfoil setup in uniform high-Reynolds number flow is investigated. The upstream airfoil is considered rigid and is actuated at its leading edge with small-amplitude harmonic pitching motion. The downstream airfoil is taken passive and elastic, with its motion forced by the vortex-street excitation of the upstream airfoil. The non-linear near-field description is obtained via potential thin-airfoil theory. It is then applied as a source term into the Powell-Howe acoustic analogy to yield the far-field dipole radiation of the system. To assess the effect of downstream-airfoil elasticity, results are compared with counterpart calculations for a non-elastic setup, where the downstream airfoil is rigid and stationary. Depending on the separation distance between airfoils, airfoil-motion and airfoil-wake dynamics shift between in-phase (synchronized) and counter-phase behaviors. Consequently, downstream airfoil elasticity may act to amplify or suppress sound through the direct contribution of elastic-airfoil motion to the total signal. Resonance-type motion of the elastic airfoil is found when the upstream airfoil is actuated at the least stable eigenfrequency of the downstream structure. This, again, results in system sound amplification or suppression, depending on the separation distance between airfoils. With increasing actuation frequency, the acoustic signal becomes dominated by the direct contribution of the upstream airfoil motion, whereas the relative contribution of the elastic airfoil to the total signature turns negligible.
A comparison of design variables for control theory based airfoil optimization
NASA Technical Reports Server (NTRS)
Reuther, James; Jameson, Antony
1995-01-01
This paper describes the implementation of optimization techniques based on control theory for airfoil design. In our previous work in the area it was shown that control theory could be employed to devise effective optimization procedures for two-dimensional profiles by using either the potential flow or the Euler equations with either a conformal mapping or a general coordinate system. We have also explored three-dimensional extensions of these formulations recently. The goal of our present work is to demonstrate the versatility of the control theory approach by designing airfoils using both Hicks-Henne functions and B-spline control points as design variables. The research also demonstrates that the parameterization of the design space is an open question in aerodynamic design.
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.
Airfoil flutter model suspension system
NASA Technical Reports Server (NTRS)
Reed, Wilmer H. (Inventor)
1987-01-01
A wind tunnel suspension system for testing flutter models under various loads and at various angles of attack is described. The invention comprises a mounting bracket assembly affixing the suspension system to the wind tunnel, a drag-link assembly and a compound spring arrangement comprises a plunge spring working in opposition to a compressive spring so as to provide a high stiffness to trim out steady state loads and simultaneously a low stiffness to dynamic loads. By this arrangement an airfoil may be tested for oscillatory response in both plunge and pitch modes while being held under high lifting loads in a wind tunnel.
Design optimization of transonic airfoils
NASA Technical Reports Server (NTRS)
Joh, C.-Y.; Grossman, B.; Haftka, R. T.
1991-01-01
Numerical optimization procedures were considered for the design of airfoils in transonic flow based on the transonic small disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented with an accurate approximation of the wave drag based on the Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without remeshing the grid. Two effective design procedures producing converged designs in approximately 10 global iterations were developed: interchanging the role of the objective function and constraint and the direct lift maximization with move limits which were fixed absolute values of the design variables.
Leading edge embedded fan airfoil concept -- A new powered high lift technology
NASA Astrophysics Data System (ADS)
Phan, Nhan Huu
A new powered-lift airfoil concept called Leading Edge Embedded Fan (LEEF) is proposed for Extremely Short Take-Off and Landing (ESTOL) and Vertical Take-Off and Landing (VTOL) applications. The LEEF airfoil concept is a powered-lift airfoil concept capable of generating thrust and very high lift-coefficient at extreme angles-of attack (AoA). It is designed to activate only at the take-off and landing phases, similar to conventional flaps or slats, allowing the aircraft to operate efficiently at cruise in its conventional configuration. The LEEF concept consists of placing a crossflow fan (CFF) along the leading-edge (LE) of the wing, and the housing is designed to alter the airfoil shape between take-off/landing and cruise configurations with ease. The unique rectangular cross section of the crossflow fan allows for its ease of integration into a conventional subsonic wing. This technology is developed for ESTOL aircraft applications and is most effectively applied to General Aviation (GA) aircraft. Another potential area of application for LEEF is tiltrotor aircraft. Unlike existing powered high-lift systems, the LEEF airfoil uses a local high-pressure air source from cross-flow fans, does not require ducting, and is able to be deployed using distributed electric power systems throughout the wing. In addition to distributed lift augmentation, the LEEF system can provide additional thrust during takeoff and landing operation to supplement the primary cruise propulsion system. Two-dimensional (2D) and three-dimensional (3D) Computational Fluid Dynamics (CFD) simulations of a conventional airfoil/wing using the NACA 63-3-418 section, commonly used in GA, and a LEEF airfoil/wing embedded into the same airfoil section were carried out to evaluate the advantages of and the costs associated with implementing the LEEF concept. Computational results show that significant lift and augmented thrust are available during LEEF operation while requiring only moderate fan power
NASA low- and medium-speed airfoil development
NASA Technical Reports Server (NTRS)
Mcghee, R. J.; Beasley, W. D.; Whitcomb, R. T.
1979-01-01
The status of NASA low and medium speed airfoil research is discussed. Effects of airfoil thickness-chord ratios varying from 9 percent to 21 percent on the section characteristics for a design lift coefficient of 0.40 are presented for the initial low speed family of airfoils. Also, modifications to the 17-percent low-speed airfoil to reduce the pitching-moment coefficient and to the 21-percent low speed airfoil results are shown for two new medium speed airfoils with thickness ratios of 13 percent and 17 percent and design-lift coefficients of 0.30. Applications of NASA-developed airfoils to general aviation aircraft are summarized.
Three-dimensional aerodynamic shape optimization using discrete sensitivity analysis
NASA Technical Reports Server (NTRS)
Burgreen, Gregory W.
1995-01-01
An aerodynamic shape optimization procedure based on discrete sensitivity analysis is extended to treat three-dimensional geometries. The function of sensitivity analysis is to directly couple computational fluid dynamics (CFD) with numerical optimization techniques, which facilitates the construction of efficient direct-design methods. The development of a practical three-dimensional design procedures entails many challenges, such as: (1) the demand for significant efficiency improvements over current design methods; (2) a general and flexible three-dimensional surface representation; and (3) the efficient solution of very large systems of linear algebraic equations. It is demonstrated that each of these challenges is overcome by: (1) employing fully implicit (Newton) methods for the CFD analyses; (2) adopting a Bezier-Bernstein polynomial parameterization of two- and three-dimensional surfaces; and (3) using preconditioned conjugate gradient-like linear system solvers. Whereas each of these extensions independently yields an improvement in computational efficiency, the combined effect of implementing all the extensions simultaneously results in a significant factor of 50 decrease in computational time and a factor of eight reduction in memory over the most efficient design strategies in current use. The new aerodynamic shape optimization procedure is demonstrated in the design of both two- and three-dimensional inviscid aerodynamic problems including a two-dimensional supersonic internal/external nozzle, two-dimensional transonic airfoils (resulting in supercritical shapes), three-dimensional transport wings, and three-dimensional supersonic delta wings. Each design application results in realistic and useful optimized shapes.
The S407, S409, and S410 Airfoils
2010-08-01
coefficient at point B is not as low as at point A, unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is...feature results in a leading-edge shape that produces a suction peak at higher lift coefficients, which ensures that transition on the upper surface...like sketch 3. Sketch 3 No suction peak exists at the leading edge. Instead, a rounded peak occurs aft of the leading edge, which allows some laminar
A Rapid Distortion Theory modified turbulence spectra for semi-analytical airfoil noise prediction
NASA Astrophysics Data System (ADS)
Santana, Leandro D.; Christophe, Julien; Schram, Christophe; Desmet, Wim
2016-11-01
This paper proposes an implementation of the Rapid Distortion Theory, for the prediction of the noise resulting from the interaction of an airfoil with incoming turbulence. In the framework of the semi-analytical modeling strategy known as Amiet's theory, this interaction mechanism is treated in a linearized form where the airfoil thickness, camber and angle of attack are assumed negligible, leading to a frozen turbulence description of the incident gust. Important semi-analytical developments have been proposed in the literature to improve the modeling of the gust-airfoil interaction accounting for parallel and skewed gusts, non-rectangular linearized airfoil shapes or blade tip effects. This work is rather focused on the investigation of the distortion of turbulence that occurs in the vicinity of the airfoil leading edge, compared with Rapid Distortion Theory, where main results are briefly reminded in this paper. The main contribution of this work is a detailed experimental investigation of the evolution of turbulent quantities relevant to noise production, performed in the close vicinity of the airfoil leading edge subjected to grid turbulence, by means of stereoscopic Particle Image Velocimetry measurements. The results indicate that the distortion effects are concentrated in a narrow region close to the stagnation point of the leading edge, with dimension of the order of its radius of curvature. Additionally, it is shown that the turbulence intensity grows significantly as the flow approaches the airfoil leading-edge. Based on those results, a modified turbulence spectrum is proposed to describe the incoming turbulence in Amiet's theory. The sound predictions show a significantly better match with acoustic measurements than using the original turbulence model.
In-flight measurement of ice growth on an airfoil using an array of ultrasonic transducers
NASA Technical Reports Server (NTRS)
Hansman, R. John, Jr.; Kirby, Mark S.; Mcknight, Robert C.; Humes, Robert L.
1988-01-01
Results of preliminary tests to measure ice growth on an airfoil during flight icing conditions are presented. Ultrasonic pulse echo measurements of ice thickness are obtained from an array of eight ultrasonic transducers mounted flush with the leading edge of the airfoil. These thickness measurements are used to document the evolution of the ice shape during the encounter in the form of successive ice profiles. Results from 3 research flights are presented and discussed. The accuracy of the ultrasonic measurements is found to be within 0.5 mm of mechanical and stereo photograph measurements of the ice accretion.
Effectiveness of spoilers on the GA(W)-1 airfoil with a high performance Fowler flap
NASA Technical Reports Server (NTRS)
Wentz, W. H., Jr.
1975-01-01
Two-dimensional wind-tunnel tests were conducted to determine effectiveness of spoilers applied to the GA(W)-1 airfoil. Tests of several spoiler configurations show adequate control effectiveness with flap nested. It is found that providing a vent path allowing lower surface air to escape to the upper surface as the spoiler opens alleviates control reversal and hysteresis tendencies. Spoiler cross-sectional shape variations generally have a modest influence on control characteristics. A series of comparative tests of vortex generators applied to the (GA-W)-1 airfoil show that triangular planform vortex generators are superior to square planform vortex generators of the same span.
On the plane potential flow past a lattice of arbitrary airfoils
NASA Technical Reports Server (NTRS)
Garrick, I E
1944-01-01
The two-dimensional, incompressible potential flow past a lattice of airfoils of arbitrary shape is investigated theoretically. The problem is treated by usual methods of conformal mapping in several stages, one stage corresponding to the mapping of the framework of the arbitrary line lattice and another significant stage corresponding to the Theodorsen method for the mapping of the arbitrary single wing profile into a circle. A particular feature in the theoretical treatment is the special handling of the regions at an infinite distance in front of and behind the lattice. Expressions are given for evaluation of the velocity and pressure distribution at the airfoil boundary. An illustrative numerical example is included.
Parameterization of solar cells
NASA Technical Reports Server (NTRS)
Appelbaum, J.; Chait, A.; Thompson, D.
1992-01-01
The aggregation (sorting) of the individual solar cells into an array is commonly based on a single operating point on the current-voltage (I-V) characteristic curve. An alternative approach for cell performance prediction and cell screening is provided by modeling the cell using an equivalent electrical circuit, in which the parameters involved are related to the physical phenomena in the device. These analytical models may be represented by a double exponential I-V characteristic with seven parameters, by a double exponential model with five parameters, or by a single exponential equation with four or five parameters. In this article we address issues concerning methodologies for the determination of solar cell parameters based on measured data points of the I-V characteristic, and introduce a procedure for screening of solar cells for arrays. We show that common curve fitting techniques, e.g., least squares, may produce many combinations of parameter values while maintaining a good fit between the fitted and measured I-V characteristics of the cell. Therefore, techniques relying on curve fitting criteria alone cannot be directly used for cell parameterization. We propose a consistent procedure which takes into account the entire set of parameter values for a batch of cells. This procedure is based on a definition of a mean cell representing the batch, and takes into account the relative contribution of each parameter to the overall goodness of fit. The procedure is demonstrated on a batch of 50 silicon cells for Space Station Freedom.
Ice Accretions and Icing Effects for Modern Airfoils
NASA Technical Reports Server (NTRS)
Addy, Harold E., Jr.
2000-01-01
Icing tests were conducted to document ice shapes formed on three different two-dimensional airfoils and to study the effects of the accreted ice on aerodynamic performance. The models tested were representative of airfoil designs in current use for each of the commercial transport, business jet, and general aviation categories of aircraft. The models were subjected to a range of icing conditions in an icing wind tunnel. The conditions were selected primarily from the Federal Aviation Administration's Federal Aviation Regulations 25 Appendix C atmospheric icing conditions. A few large droplet icing conditions were included. To verify the aerodynamic performance measurements, molds were made of selected ice shapes formed in the icing tunnel. Castings of the ice were made from the molds and placed on a model in a dry, low-turbulence wind tunnel where precision aerodynamic performance measurements were made. Documentation of all the ice shapes and the aerodynamic performance measurements made during the icing tunnel tests is included in this report. Results from the dry, low-turbulence wind tunnel tests are also presented.
Trailing edge modifications for flatback airfoils.
Kahn, Daniel L.; van Dam, C.P.; Berg, Dale E.
2008-03-01
The adoption of blunt trailing edge airfoils (also called flatback airfoils) for the inboard region of large wind turbine blades has been proposed. Blunt trailing edge airfoils would not only provide a number of structural benefits, such as increased structural volume and ease of fabrication and handling, but they have also been found to improve the lift characteristics of thick airfoils. Therefore, the incorporation of blunt trailing edge airfoils would allow blade designers to more freely address the structural demands without having to sacrifice aerodynamic performance. These airfoils do have the disadvantage of generating high levels of drag as a result of the low-pressure steady or periodic flow in the near-wake of the blunt trailing edge. Although for rotors, the drag penalty appears secondary to the lift enhancement produced by the blunt trailing edge, high drag levels are of concern in terms of the negative effect on the torque and power generated by the rotor. Hence, devices are sought that mitigate the drag of these airfoils. This report summarizes the literature on bluff body vortex shedding and bluff body drag reduction devices and proposes four devices for further study in the wind tunnel.
Prediction of airfoil stall using Navier-Stokes equations in streamline coordinates
NASA Technical Reports Server (NTRS)
Choi, D. H.; Sohn, C. H.; Oh, C. S.
1992-01-01
A Navier-Stokes procedure to calculate the flow about an airfoil at incidence was developed. The parabolized equations are solved in the streamline coordinates generated for an arbitrary airfoil shape using conformal mapping. A modified k-epsilon turbulence model is applied in the entire domain, but the eddy viscosity in the laminar region is suppressed artificially to simulate the region correctly. The procedure was applied to airfoils at various angles of attack, and the results are quite satisfactory for both laminar and turbulent flows. It is shown that the present choice of the coordinate system reduces the error due to numerical diffusion, and that the lift is accurately predicted for a wide range of incidence.
A Wind Tunnel Study of Icing Effects on a Business Jet Airfoil
NASA Technical Reports Server (NTRS)
Addy, Harold E., Jr.; Broeren, Andy P.; Zoeckler, Joesph G.; Lee, Sam
2003-01-01
Aerodynamic wind tunnel tests were conducted to study the effects of various ice accretions on the aerodynamic performance of a 36-inch chord, two-dimensional business jet airfoil. Eight different ice shape configurations were tested. Four were castings made from molds of ice shapes accreted in an icing wind tunnel. Two were made using computationally smoothed tracings of two of the ice shapes accreted in the icing tunnel. These smoothed profiles were then extended in the spanwise direction to form a two-dimensional ice shape. The final two configurations were formed by applying grit to the smoothed ice shapes. The ice shapes resulted in as much as 48% reduction in maximum lift coefficient from that of the clean airfoil. Large increases in drag and changes in pitching moment were also observed. The castings and their corresponding smoothed counterparts yielded similar results. Little change in performance was observed with the addition of grit to the smoothed ice shapes. Changes in the Reynolds number (from 3 x 10(exp 6) to 10.5 x 10(exp 6) and Mach number (from 0.12 to 0.28) did not significantly affect the iced-airfoil performance coefficients.
Turbine airfoil with outer wall thickness indicators
Marra, John J; James, Allister W; Merrill, Gary B
2013-08-06
A turbine airfoil usable in a turbine engine and including a depth indicator for determining outer wall blade thickness. The airfoil may include an outer wall having a plurality of grooves in the outer surface of the outer wall. The grooves may have a depth that represents a desired outer surface and wall thickness of the outer wall. The material forming an outer surface of the outer wall may be removed to be flush with an innermost point in each groove, thereby reducing the wall thickness and increasing efficiency. The plurality of grooves may be positioned in a radially outer region of the airfoil proximate to the tip.
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.
Second-stage turbine bucket airfoil
Wang, John Zhiqiang; By, Robert Romany; Sims, Calvin L.; Hyde, Susan Marie
2002-01-01
The second-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinate values defining the airfoil profile at each distance Z. The X and Y values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket. The second-stage wheel has sixty buckets.
Third-stage turbine bucket airfoil
Pirolla, Peter Paul; Siden, Gunnar Leif; Humanchuk, David John; Brassfield, Steven Robert; Wilson, Paul Stuart
2002-01-01
The third-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinates defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.
Pressure Distribution Over Airfoils with Fowler Flaps
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Anderson, Walter B
1938-01-01
Report presents the results of tests made of a Clark y airfoil with a Clark y Fowler flap and of an NACA 23012 airfoil with NACA Fowler flaps. Some of the tests were made in the 7 by 10-foot wind tunnel and others in the 5-foot vertical wind tunnel. The pressures were measured on the upper and lower surfaces at one chord section both on the main airfoils and on the flaps for several angles of attack with the flaps located at the maximum-lift settings. A test installation was used in which the model was mounted in the wind tunnel between large end planes so that two-dimensional flow was approximated. The data are given in the form of pressure-distribution diagrams and as plots of calculated coefficients for the airfoil-and-flap combinations and for the flaps alone.
Turbine airfoil to shroud attachment method
Campbell, Christian X; Kulkarni, Anand A; James, Allister W; Wessell, Brian J; Gear, Paul J
2014-12-23
Bi-casting a platform (50) onto an end portion (42) of a turbine airfoil (31) after forming a coating of a fugitive material (56) on the end portion. After bi-casting the platform, the coating is dissolved and removed to relieve differential thermal shrinkage stress between the airfoil and platform. The thickness of the coating is varied around the end portion in proportion to varying amounts of local differential process shrinkage. The coating may be sprayed (76A, 76B) onto the end portion in opposite directions parallel to a chord line (41) of the airfoil or parallel to a mid-platform length (80) of the platform to form respective layers tapering in thickness from the leading (32) and trailing (34) edges along the suction side (36) of the airfoil.
Liang, George [Palm City, FL
2011-01-18
An airfoil is provided for a gas turbine comprising an outer structure comprising a first wall, an inner structure comprising a second wall spaced relative to the first wall such that a cooling gap is defined between at least portions of the first and second walls, and seal structure provided within the cooling gap between the first and second walls for separating the cooling gap into first and second cooling fluid impingement gaps. An inner surface of the second wall may define an inner cavity. The inner structure may further comprise a separating member for separating the inner cavity of the inner structure into a cooling fluid supply cavity and a cooling fluid collector cavity. The second wall may comprise at least one first impingement passage, at least one second impingement passage, and at least one bleed passage.
The Influence of Microphysical Cloud Parameterization on Microwave Brightness Temperatures
NASA Technical Reports Server (NTRS)
Skofronick-Jackson, Gail M.; Gasiewski, Albin J.; Wang, James R.; Zukor, Dorothy J. (Technical Monitor)
2000-01-01
The microphysical parameterization of clouds and rain-cells plays a central role in atmospheric forward radiative transfer models used in calculating passive microwave brightness temperatures. The absorption and scattering properties of a hydrometeor-laden atmosphere are governed by particle phase, size distribution, aggregate density., shape, and dielectric constant. This study identifies the sensitivity of brightness temperatures with respect to the microphysical cloud parameterization. Cloud parameterizations for wideband (6-410 GHz observations of baseline brightness temperatures were studied for four evolutionary stages of an oceanic convective storm using a five-phase hydrometeor model in a planar-stratified scattering-based radiative transfer model. Five other microphysical cloud parameterizations were compared to the baseline calculations to evaluate brightness temperature sensitivity to gross changes in the hydrometeor size distributions and the ice-air-water ratios in the frozen or partly frozen phase. The comparison shows that, enlarging the rain drop size or adding water to the partly Frozen hydrometeor mix warms brightness temperatures by up to .55 K at 6 GHz. The cooling signature caused by ice scattering intensifies with increasing ice concentrations and at higher frequencies. An additional comparison to measured Convection and Moisture LA Experiment (CAMEX 3) brightness temperatures shows that in general all but, two parameterizations produce calculated T(sub B)'s that fall within the observed clear-air minima and maxima. The exceptions are for parameterizations that, enhance the scattering characteristics of frozen hydrometeors.
Transonic airfoil design using Cartesian coordinates
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1976-01-01
A numerical technique for designing transonic airfoils having a prescribed pressure distribution (the inverse problem) is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that inverse boundary conditions and Cartesian coordinates are used. The method is a direct-inverse approach that controls trailing-edge closure. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.
Transonic Airfoils with a Given Pressure Distribution,
1981-06-01
erovse sidst necesosar mod Ideatify b lock mmb)L An inverse design procedure for airfoils, based on hodograph techniques, has been developed. For...w L-:- " " -- - r- L i -- _ 9 ABSTRACT An inverse design procedure for airfoils, based on hodograph tech...generated in the hodograph plane by Nieuwand,5 Bauer, Garabedian and Korn,6 Boerstoel and Huizing,7 and Sobieczky.8 More recently, the development of
Unsteady Pressure Distributions on Airfoils in Cascade.
1980-04-01
of thin airfoil theory has been used by Henderson (-ftj’ and Bruce (1-7-)’to derive expressions for the unsteady response which includes the cascade...model in conjunction with the assumptions of thin airfoil theory has been used by Henderson (16) and Bruce (17) to derive expressions for the unsteady...effect, that is, a sharp change in the unsteady lift when the disturbance wavelength equals the blade spacing. Bruce (19) further extends this theory to
Propulsion by active and passive airfoil oscillation
NASA Astrophysics Data System (ADS)
Mackowski, A. W.; Williamson, C. H. K.
2013-11-01
Oscillating airfoils have been the subject of much research both as a mechanism of propulsion in engineering devices as well as a model of understanding how fish, birds, and insects produce thrust and maneuvering forces. Additionally, the jet or wake generated by an oscillating airfoil exhibits a multitude of vortex patterns, which are an interesting study in their own right. We present PIV measurements of the vortex flow behind an airfoil undergoing controlled pitching oscillations at moderate Reynolds number. As a method of propulsion, oscillating foils have been found to be capable performers when undergoing both pitching and heaving motions [Anderson et al. 1998]. While an airfoil undergoing only pitching motion is a relatively inefficient propulsor, we examine the effect of adding passive dynamics to the system: for example, actuated pitching with a passive spring in the heave direction. Practically speaking, a mechanical system with such an arrangement has the potential to reduce the cost and complexity of an oscillating airfoil propulsor. To study an airfoil undergoing both active and passive motion, we employ our ``cyber-physical fluid dynamics'' technique [Mackowski & Williamson, 2011] to simulate the effects of passive dynamics in a physical experiment.
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.
Simulation of a Controlled Airfoil with Jets
NASA Technical Reports Server (NTRS)
Allan, Brian G.; Holt, Maurice; Packard, Andrew
1997-01-01
Numerical simulations of a two-dimensional airfoil, controlled by an applied moment in pitch and an airfoil controlled by jets, were investigated. These simulations couple the Reynolds-averaged Navier-Stokes equations and Euler's equations of rigid body motion, with an active control system. Controllers for both systems were designed to track altitude commands and were evaluated by simulating a closed-loop altitude step response using the coupled system. The airfoil controlled by a pitching moment used an optimal state feedback controller. A closed-loop simulation, of the airfoil with an applied moment, showed that the trajectories compared very well with quasi-steady aerodynamic theory, providing a measure of validation. The airfoil with jets used a controller designed by robust control methods. A linear plant model for this system was identified using open-loop data generated by the nonlinear coupled system. A closed-loop simulation of the airfoil with jets, showed good tracking of an altitude command. This simulation also showed oscillations in the control input as a result of dynamics not accounted for in the control design. This research work demonstrates how computational fluid dynamics, coupled with rigid body dynamics, and a control law can be used to prototype control systems in problematic nonlinear flight regimes.
Isolated and cascade airfoils with prescribed velocity distribution
NASA Technical Reports Server (NTRS)
Goldstein, Arthur W; Jerison, Meyer
1947-01-01
An exact solution of the problem of designing an airfoil with a prescribed velocity distribution on the suction surface in a given uniform flow of an incompressible perfect fluid is obtained by replacing the boundary of the airfoil by vortices. By this device, a method of solution is developed that is applicable both to isolated airfoils and to airfoils in cascade. The conformal transformation of the designed airfoil into a circle can then be obtained and the velocity distribution at any angle of attack computed. Numerical illustrations of the method are given for the airfoil in cascade.
NASA Technical Reports Server (NTRS)
Kolesar, C. E.
1987-01-01
Research activity on an airfoil designed for a large airplane capable of very long endurance times at a low Mach number of 0.22 is examined. Airplane mission objectives and design optimization resulted in requirements for a very high design lift coefficient and a large amount of laminar flow at high Reynolds number to increase the lift/drag ratio and reduce the loiter lift coefficient. Natural laminar flow was selected instead of distributed mechanical suction for the measurement technique. A design lift coefficient of 1.5 was identified as the highest which could be achieved with a large extent of laminar flow. A single element airfoil was designed using an inverse boundary layer solution and inverse airfoil design computer codes to create an airfoil section that would achieve performance goals. The design process and results, including airfoil shape, pressure distributions, and aerodynamic characteristics are presented. A two dimensional wind tunnel model was constructed and tested in a NASA Low Turbulence Pressure Tunnel which enabled testing at full scale design Reynolds number. A comparison is made between theoretical and measured results to establish accuracy and quality of the airfoil design technique.
NASA Technical Reports Server (NTRS)
Allison, D. O.
1976-01-01
A 20.8 percent-thick airfoil shape was designed to have shockless inviscid flow at a Mach number of 0.68 and a lift coefficient of 0.40. In order to determine the actual airfoils which would yield this same shockless flow when viscous effects are included, boundary layer displacement thicknesses were subtracted from the inviscid shape for Reynolds numbers of 100 and 35 million. This process yielded airfoils with thicknesses of 20.7 and 20.6 percent, respectively. Subtraction of boundary layer displacement thicknesses for Reynolds numbers below 35 million yielded nonphysical airfoils, that is airfoils with negative thicknesses near tHe trailing edge. The pitching moment about the quarter-chord point at the design condition was -0.082 for the inviscid shape and, consequently, for both airfoils. Off-design calculations for the two airfoils were made using a computer program which provides for the interaction of the inviscid flow and boundary layer solutions. The pressure distributions of the airfoils were shockless for conditions from the design point to lower Mach numbers and lift coefficients. No boundary layer separation was predicted except in the last 3 percent chord on the upper surface.
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.
Parameter study of simplified dragonfly airfoil geometry at Reynolds number of 6000.
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
Transonic Aerodynamic Characteristics of Two Wedge Airfoil Sections Including Unsteady Flow Studies
NASA Technical Reports Server (NTRS)
Johnston, Patrick J.
1959-01-01
A two-dimensional wind-tunnel investigation has been conducted on a 20-percent-thick single-wedge airfoil section. Steady-state forces and moments were determined from pressure measurements at Mach numbers from 0.70 to about 1.25. Additional information on the flows about the single wedge is provided by means of instantaneous pressure measurements at Mach numbers up to unity. Pressure distributions were also obtained on a symmetrical double-wedge or diamond-shaped profile which had the same leading-edge included angle as the single-wedge airfoil. A comparison of the data on the two profiles to provide information on the effects of the afterbody showed that with the exception of drag, the single-wedge profile proved to be aerodynamically superior to the diamond profile in all respects. The lift effectiveness of the single-wedge airfoil section far exceeded that of conventional thin airfoil sections over the speed range of the investigation. Pitching-moment irregularities, caused by negative loadings near the trailing edge, generally associated with conventional airfoils of equivalent thicknesses were not exhibited by the single-wedge profile. Moderately high pulsating pressures existing over the base of the single-wedge airfoil section were significantly reduced as the Mach number was increased beyond 0.92 and the boundaries of the dead airspace at the base of the model converged to eliminate the vortex street in the wake. Increasing the leading-edge radius from 0 to 1 percent of the chord had a minor effect on the steady-state forces and generally raised the level of pressure pulsations over the forward part of the single-wedge profile.
NASA Technical Reports Server (NTRS)
Wood, Richard M.; Byrd, James E.; Wesselmann, Gary F.
1992-01-01
An assessment of the influence of airfoil geometry on delta wing leading edge vortex flow and vortex induced aerodynamics at supersonic speeds is discussed. A series of delta wing wind tunnel models were tested over a Mach number range from 1.7 to 2.0. The model geometric variables included leading edge sweep and airfoil shape. Surface pressure data, vapor screen, and oil flow photograph data were taken to evaluate the complex structure of the vortices and shocks on the family of wings tested. The data show that airfoil shape has a significant impact on the wing upper surface flow structure and pressure distribution, but has a minimal impact on the integrated upper surface pressure increments.
Program manual for the Eppler airfoil inversion program
NASA Technical Reports Server (NTRS)
Thomson, W. G.
1975-01-01
A computer program is described for calculating the profile of an airfoil as well as the boundary layer momentum thickness and energy form parameter. The theory underlying the airfoil inversion technique developed by Eppler is discussed.
Summary of Cumulus Parameterization Workshop
NASA Technical Reports Server (NTRS)
Tao, Wei-Kuo; Starr, David OC.; Hou, Arthur; Newman, Paul; Sud, Yogesh
2002-01-01
A workshop on cumulus parameterization took place at the NASA Goddard Space Flight Center from December 3-5, 2001. The major objectives of this workshop were (1) to review the problem of representation of moist processes in large-scale models (mesoscale models, Numerical Weather Prediction models and Atmospheric General Circulation Models), (2) to review the state-of-the-art in cumulus parameterization schemes, and (3) to discuss the need for future research and applications. There were a total of 31 presentations and about 100 participants from the United States, Japan, the United Kingdom, France and South Korea. The specific presentations and discussions during the workshop are summarized in this paper.
AirfoilPrep.py Documentation: Release 0.1.0
Ning, S. A.
2013-09-01
AirfoilPrep.py provides functionality to preprocess aerodynamic airfoil data. Essentially, the module is an object oriented version of the AirfoilPrep spreadsheet with additional functionality and is written in the Python language. It allows the user to read in two-dimensional aerodynamic airfoil data, apply three-dimensional rotation corrections for wind turbine applications, and extend the datato very large angles of attack. This document discusses installation, usage, and documentation of the module.
Critical Mach Numbers of Thin Airfoil Sections with Plain Flaps
NASA Technical Reports Server (NTRS)
Pardee, Otway O'm.; Heaslet, Max A.
1946-01-01
Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils.
Geometries for roughness shapes in laminar flow
NASA Technical Reports Server (NTRS)
Holmes, Bruce J. (Inventor); Martin, Glenn L. (Inventor); Domack, Christopher S. (Inventor); Obara, Clifford J. (Inventor); Hassan, Ahmed A. (Inventor)
1986-01-01
A passive interface mechanism between upper and lower skin structures, and a leading edge structure of a laminar flow airfoil is described. The interface mechanism takes many shapes. All are designed to be different than the sharp orthogonal arrangement prevalent in the prior art. The shapes of the interface structures are generally of two types: steps away from the centerline of the airfoil with a sloping surface directed toward the trailing edge and, the other design has a gap before the sloping surface. By properly shaping the step, the critical step height is increased by more than 50% over the orthogonal edged step.
Coating-Substrate Systems for Thermomechanically Durable Turbine Airfoils
2015-06-30
Technical Report 4. TITLE AND SUBTITLE Coating - Substrate Systems for Thermomechanically Durable Turbine Airfoils 6. AUTHOR(S) Dr. Tresa Pollock 3...Thermomechanically Durable Turbine Airfoils Final Report ONRGrant#N00014-l 1-1-0616 Technical Contact (Principal Investigator) Tresa M. Pollock Materials...Substrate Systems for Thermomechanically Durable Turbine Airfoils 1. Summary In the severe operating environments encountered in Naval ship
Computing Aerodynamic Performance of a 2D Iced Airfoil: Blocking Topology and Grid Generation
NASA Technical Reports Server (NTRS)
Chi, X.; Zhu, B.; Shih, T. I.-P.; Slater, J. W.; Addy, H. E.; Choo, Yung K.; Lee, Chi-Ming (Technical Monitor)
2002-01-01
The ice accrued on airfoils can have enormously complicated shapes with multiple protruded horns and feathers. In this paper, several blocking topologies are proposed and evaluated on their ability to produce high-quality structured multi-block grid systems. A transition layer grid is introduced to ensure that jaggedness on the ice-surface geometry do not to propagate into the domain. This is important for grid-generation methods based on hyperbolic PDEs (Partial Differential Equations) and algebraic transfinite interpolation. A 'thick' wrap-around grid is introduced to ensure that grid lines clustered next to solid walls do not propagate as streaks of tightly packed grid lines into the interior of the domain along block boundaries. For ice shapes that are not too complicated, a method is presented for generating high-quality single-block grids. To demonstrate the usefulness of the methods developed, grids and CFD solutions were generated for two iced airfoils: the NLF0414 airfoil with and without the 623-ice shape and the B575/767 airfoil with and without the 145m-ice shape. To validate the computations, the computed lift coefficients as a function of angle of attack were compared with available experimental data. The ice shapes and the blocking topologies were prepared by NASA Glenn's SmaggIce software. The grid systems were generated by using a four-boundary method based on Hermite interpolation with controls on clustering, orthogonality next to walls, and C continuity across block boundaries. The flow was modeled by the ensemble-averaged compressible Navier-Stokes equations, closed by the shear-stress transport turbulence model in which the integration is to the wall. All solutions were generated by using the NPARC WIND code.
Evaluation of Icing Scaling on Swept NACA 0012 Airfoil Models
NASA Technical Reports Server (NTRS)
Tsao, Jen-Ching; Lee, Sam
2012-01-01
Icing scaling tests in the NASA Glenn Icing Research Tunnel (IRT) were performed on swept wing models using existing recommended scaling methods that were originally developed for straight wing. Some needed modifications on the stagnation-point local collection efficiency (i.e., beta(sub 0) calculation and the corresponding convective heat transfer coefficient for swept NACA 0012 airfoil models have been studied and reported in 2009, and the correlations will be used in the current study. The reference tests used a 91.4-cm chord, 152.4-cm span, adjustable sweep airfoil model of NACA 0012 profile at velocities of 100 and 150 knot and MVD of 44 and 93 mm. Scale-to-reference model size ratio was 1:2.4. All tests were conducted at 0deg angle of attack (AoA) and 45deg sweep angle. Ice shape comparison results were presented for stagnation-point freezing fractions in the range of 0.4 to 1.0. Preliminary results showed that good scaling was achieved for the conditions test by using the modified scaling methods developed for swept wing icing.
Aerodynamic Control of a Pitching Airfoil using Distributed Active Bleed
NASA Astrophysics Data System (ADS)
Kearney, John; Glezer, Ari
2012-11-01
Aero-effected flight control using distributed active bleed driven by pressure differences across lifting surface and regulated by integrated louver actuators is investigated in wind tunnel experiments. The interaction between unsteady bleed and the cross flows alters the apparent aerodynamic shape of the lifting surface by regulating the accumulation and shedding of vorticity concentrations, and consequently the distributions of forces and moments. The present experiments are conducted using a 2-D dynamically-pitching VR-7 airfoil model from pre- to post-stall angles of attack. The effects of leading edge bleed at high angles of attack on the formation and evolution of the dynamic stall vorticity concentrations are investigated at high reduced frequencies (k > 0.1) using PIV phase-locked to the airfoil's motion. The time-dependent bleed enables broad-range variation in lift and pitching moment with significant extension of the stall margin. In particular, bleed actuation reduces the extent of ``negative damping'' or pitching moment instability with minimal lift penalty. Supported by NTRC-VLRCOE, monitored by Dr. Mike Rutkowski.
Meshless thin-shell simulation based on global conformal parameterization.
Guo, Xiaohu; Li, Xin; Bao, Yunfan; Gu, Xianfeng; Qin, Hong
2006-01-01
This paper presents a new approach to the physically-based thin-shell simulation of point-sampled geometry via explicit, global conformal point-surface parameterization and meshless dynamics. The point-based global parameterization is founded upon the rigorous mathematics of Riemann surface theory and Hodge theory. The parameterization is globally conformal everywhere except for a minimum number of zero points. Within our parameterization framework, any well-sampled point surface is functionally equivalent to a manifold, enabling popular and powerful surface-based modeling and physically-based simulation tools to be readily adapted for point geometry processing and animation. In addition, we propose a meshless surface computational paradigm in which the partial differential equations (for dynamic physical simulation) can be applied and solved directly over point samples via Moving Least Squares (MLS) shape functions defined on the global parametric domain without explicit connectivity information. The global conformal parameterization provides a common domain to facilitate accurate meshless simulation and efficient discontinuity modeling for complex branching cracks. Through our experiments on thin-shell elastic deformation and fracture simulation, we demonstrate that our integrative method is very natural, and that it has great potential to further broaden the application scope of point-sampled geometry in graphics and relevant fields.
Compressor airfoil tip clearance optimization system
Little, David A.; Pu, Zhengxiang
2015-08-18
A compressor airfoil tip clearance optimization system for reducing a gap between a tip of a compressor airfoil and a radially adjacent component of a turbine engine is disclosed. The turbine engine may include ID and OD flowpath boundaries configured to minimize compressor airfoil tip clearances during turbine engine operation in cooperation with one or more clearance reduction systems that are configured to move the rotor assembly axially to reduce tip clearance. The configurations of the ID and OD flowpath boundaries enhance the effectiveness of the axial movement of the rotor assembly, which includes movement of the ID flowpath boundary. During operation of the turbine engine, the rotor assembly may be moved axially to increase the efficiency of the turbine engine.
Turbine airfoil fabricated from tapered extrusions
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.
Comparative Study of Airfoil Flow Separation Criteria
NASA Astrophysics Data System (ADS)
Laws, Nick; Kahouli, Waad; Epps, Brenden
2015-11-01
Airfoil flow separation impacts a multitude of applications including turbomachinery, wind turbines, and bio-inspired micro-aerial vehicles. In order to achieve maximum performance, some devices operate near the edge of flow separation, and others use dynamic flow separation advantageously. Numerous criteria exist for predicting the onset of airfoil flow separation. This talk presents a comparative study of a number of such criteria, with emphasis paid to speed and accuracy of the calculations. We evaluate the criteria using a two-dimensional unsteady vortex lattice method, which allows for rapid analysis (on the order of seconds instead of days for a full Navier-Stokes solution) and design of optimal airfoil geometry and kinematics. Furthermore, dynamic analyses permit evaluation of dynamic stall conditions for enhanced lift via leading edge vortex shedding, commonly present in small flapping-wing flyers such as the bumblebee and hummingbird.
Vortex noise from nonrotating cylinders and airfoils
NASA Technical Reports Server (NTRS)
Schlinker, R. H.; Amiet, R. K.; Fink, M. R.
1976-01-01
An experimental study of vortex-shedding noise was conducted in an acoustic research tunnel over a Reynolds-number range applicable to full-scale helicopter tail-rotor blades. Two-dimensional tapered-chord nonrotating models were tested to simulate the effect of spanwise frequency variation on the vortex-shedding mechanism. Both a tapered circular cylinder and tapered airfoils were investigated. The results were compared with data for constant-diameter cylinder and constant-chord airfoil models also tested during this study. Far-field noise, surface pressure fluctuations, and spanwise correlation lengths were measured for each configuration. Vortex-shedding noise for tapered cylinders and airfoils was found to contain many narrowband-random peaks which occurred within a range of frequencies corresponding to a predictable Strouhal number referenced to the maximum and minimum chord. The noise was observed to depend on surface roughness and Reynolds number.
Near-wall serpentine cooled turbine airfoil
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.
CFD aerodynamic analysis of non-conventional airfoil sections for very large rotor blades
NASA Astrophysics Data System (ADS)
Papadakis, G.; Voutsinas, S.; Sieros, G.; Chaviaropoulos, T.
2014-12-01
The aerodynamic performance of flat-back and elliptically shaped airfoils is analyzed on the basis of CFD simulations. Incompressible and low-Mach preconditioned compressible unsteady simulations have been carried out using the k-w SST and the Spalart Allmaras turbulence models. Time averaged lift and drag coefficients are compared to wind tunnel data for the FB 3500-1750 flat back airfoil while amplitudes and frequencies are also recorded. Prior to separation averaged lift is well predicted while drag is overestimated keeping however the trend in the tests. The CFD models considered, predict separation with a 5° delay which is reflected on the load results. Similar results are provided for a modified NACA0035 with a rounded (elliptically shaped) trailing edge. Finally as regards the dynamic characteristics in the load signals, there is fair agreement in terms of Str number but significant differences in terms of lift and drag amplitudes.
A Genus Oblivious Approach to Cross Parameterization
Bennett, J C; Pascucci, V; Joy, K I
2008-06-16
In this paper we present a robust approach to construct a map between two triangulated meshes, M and M{prime} of arbitrary and possibly unequal genus. We introduce a novel initial alignment scheme that allows the user to identify 'landmark tunnels' and/or a 'constrained silhouette' in addition to the standard landmark vertices. To describe the evolution of non-landmark tunnels we automatically derive a continuous deformation from M to M{prime} using a variational implicit approach. Overall, we achieve a cross parameterization scheme that is provably robust in the sense that it can map M to M{prime} without constraints on their relative genus. We provide a number of examples to demonstrate the practical effectiveness of our scheme between meshes of different genus and shape.
The semi-discrete Galerkin finite element modelling of compressible viscous flow past an airfoil
NASA Technical Reports Server (NTRS)
Meade, Andrew J., Jr.
1992-01-01
A method is developed to solve the two-dimensional, steady, compressible, turbulent boundary-layer equations and is coupled to an existing Euler solver for attached transonic airfoil analysis problems. The boundary-layer formulation utilizes the semi-discrete Galerkin (SDG) method to model the spatial variable normal to the surface with linear finite elements and the time-like variable with finite differences. A Dorodnitsyn transformed system of equations is used to bound the infinite spatial domain thereby permitting the use of a uniform finite element grid which provides high resolution near the wall and automatically follows boundary-layer growth. The second-order accurate Crank-Nicholson scheme is applied along with a linearization method to take advantage of the parabolic nature of the boundary-layer equations and generate a non-iterative marching routine. The SDG code can be applied to any smoothly-connected airfoil shape without modification and can be coupled to any inviscid flow solver. In this analysis, a direct viscous-inviscid interaction is accomplished between the Euler and boundary-layer codes, through the application of a transpiration velocity boundary condition. Results are presented for compressible turbulent flow past NACA 0012 and RAE 2822 airfoils at various freestream Mach numbers, Reynolds numbers, and angles of attack. All results show good agreement with experiment, and the coupled code proved to be a computationally-efficient and accurate airfoil analysis tool.
Experiments on a Steady Low Reynolds Number Airfoil in a Shear Flow
NASA Astrophysics Data System (ADS)
Olson, David; Naguib, Ahmed; Koochesfahani, Manoochehr
2016-11-01
The aerodynamics of steady airfoils in uniform flow have received considerably more attention than that of an airfoil operating in a non-uniform flow. Inviscid theory by Tsien (1943) shows that an airfoil experiences a decrease in the zero lift angle of attack for a shear flow with uniform clockwise vorticity. The current work utilizes a shaped honeycomb technique to create a velocity profile with a large region of uniform shear in a water tunnel. Direct force measurements are implemented and validated using experiments on a circular cylinder and NACA 0012 in a uniform cross-flow. Results for a NACA 0012 airfoil with a chord Reynolds number of 1.2 ×104 in a non-uniform approach flow are compared to concurrent CFD calculations (presented in a companion talk) showing an increase in the zero lift angle of attack; in contradiction with inviscid theory. The effect of shear on the mean lift coefficient over a wide range of angles of attack is also explored. This work was supported by AFOSR Award Number FA9550-15-1-0224.
Blowing Circulation Control on a Seaplane Airfoil
NASA Astrophysics Data System (ADS)
Guo, B. D.; Liu, P. Q.; Qu, Q. L.
2011-09-01
RANS simulations are presented for blowing circulation control on a seaplane airfoil. Realizable k-epsilon turbulent model and pressure-based coupled algorithm with second-order discretization were adopted to simulate the compressible flow. Both clear and simple flap configuration were simulated with blowing momentum coefficient Cμ = 0, 0.15 and 0.30. The results show that blowing near the airfoil trailing edge could enhance the Coanda effect, delay the flow separation, and increase the lift coefficient dramatically. The blowing circulation control is promising to apply to taking off and landing of an amphibious aircraft or seaplane.
An efficient algorithm for numerical airfoil optimization
NASA Technical Reports Server (NTRS)
Vanderplaats, G. N.
1979-01-01
A new optimization algorithm is presented. The method is based on sequential application of a second-order Taylor's series approximation to the airfoil characteristics. Compared to previous methods, design efficiency improvements of more than a factor of 2 are demonstrated. If multiple optimizations are performed, the efficiency improvements are more dramatic due to the ability of the technique to utilize existing data. The method is demonstrated by application to subsonic and transonic airfoil design but is a general optimization technique and is not limited to a particular application or aerodynamic analysis.
Multi-pass cooling for turbine airfoils
Liang, George
2011-06-28
An airfoil for a turbine vane of a gas turbine engine. The airfoil includes an outer wall having pressure and suction sides, and a radially extending cooling cavity located between the pressure and suction sides. A plurality of partitions extend radially through the cooling cavity to define a plurality of interconnected cooling channels located at successive chordal locations through the cooling cavity. The cooling channels define a serpentine flow path extending in the chordal direction. Further, the cooling channels include a plurality of interconnected chambers and the chambers define a serpentine path extending in the radial direction within the serpentine path extending in the chordal direction.
Transonic flow theory of airfoils and wings
NASA Technical Reports Server (NTRS)
Garabedian, P. R.
1976-01-01
There are plans to use the supercritical wing on the next generation of commercial aircraft so as to economize on fuel consumption by reducing drag. Computer codes have served well in meeting the consequent demand for new wing sections. The possibility of replacing wind tunnel tests by computational fluid dynamics is discussed. Another approach to the supercritical wing is through shockless airfoils. A novel boundary value problem in the hodograph plane is studied that enables one to design a shockless airfoil so that its pressure distribution very nearly takes on data that are prescribed.
Advanced technology airfoil research, volume 1, part 2
NASA Technical Reports Server (NTRS)
1978-01-01
This compilation contains papers presented at the NASA Conference on Advanced Technology Airfoil Research held at Langley Research Center on March 7-9, 1978, which have unlimited distribution. This conference provided a comprehensive review of all NASA airfoil research, conducted in-house and under grant and contract. A broad spectrum of airfoil research outside of NASA was also reviewed. The major thrust of the technical sessions were in three areas: development of computational aerodynamic codes for airfoil analysis and design, development of experimental facilities and test techniques, and all types of airfoil applications.
Multiple piece turbine engine airfoil with a structural spar
Vance, Steven J.
2011-10-11
A multiple piece turbine airfoil having an outer shell with an airfoil tip that is attached to a root with an internal structural spar is disclosed. The root may be formed from first and second sections that include an internal cavity configured to receive and secure the one or more components forming the generally elongated airfoil. The internal structural spar may be attached to an airfoil tip and place the generally elongated airfoil in compression. The configuration enables each component to be formed from different materials to reduce the cost of the materials and to optimize the choice of material for each component.
Numerical solution of the Navier-Stokes equations for arbitrary 2-dimensional multi-element airfoils
NASA Technical Reports Server (NTRS)
Thompson, J. F.
1983-01-01
Numerical solutions of the Navier-Stokes equations, with an algebraic turbulence model, for time-dependent two dimensional flow about multi-element airfoils were developed. Fundamental to these solutions was the use of numerically-generated boundary-conforming curvilinear coordinate systems to allow bodies of arbitrary shape to be treated. A general two dimensional grid generation code for multiple-body configuration was written as a part of this project and made available through the COSMIC code library.
1982-07-01
Aeronautics and United States Army Space Administration Aviation Research and Ames Remrch Cente Development Command Moffett Field. California 94035 St...appear to be more important than airfoil shape in determining the dynamic- stall airloads. 1. INTRODUCTION Retreating- blade stall limits the high-speed...12.2% Thick R.A.E. Aerofoil Section. RAE Technical Report 68303, Royal Aircraft Establishment, Farnborough Hants, England, Jan. 1969. 14. Fromme, J. A
On the acoustic radiation of a pitching airfoil
NASA Astrophysics Data System (ADS)
Manela, A.
2013-07-01
We examine the acoustic far field of a thin elastic airfoil, immersed in low-Mach non-uniform stream flow, and actuated by small-amplitude sinusoidal pitching motion. The near-field fluid-structure interaction problem is analyzed using potential thin-airfoil theory, combined with a discrete vortex model to describe the evolution of airfoil trailing edge wake. The leading order dipole-sound signature of the system is investigated using Powell-Howe acoustic analogy. Compared with a pitching rigid airfoil, the results demonstrate a two-fold effect of structure elasticity on airfoil acoustic field: at actuation frequencies close to the system least stable eigenfrequency, elasticity amplifies airfoil motion amplitude and associated sound levels; however, at frequencies distant from this eigenfrequency, structure elasticity acts to absorb system kinetic energy and reduce acoustic radiation. In the latter case, and with increasing pitching frequency ωp, a rigid-airfoil setup becomes significantly noisier than an elastic airfoil, owing to an ω _p^{5/2} increase of its direct motion noise component. Unlike rigid airfoil signature, it is shown that wake sound contribution to elastic airfoil radiation is significant for all ωp. Remarkably, this contribution contains, in addition to the fundamental pitching frequency, its odd multiple harmonics, which result from nonlinear interactions between the airfoil and the wake. The results suggest that structure elasticity may serve as a viable means for design of flapping flight noise control methodologies.
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.
Navier-Stokes analysis of airfoils with leading edge ice accretions
NASA Technical Reports Server (NTRS)
Potapczuk, Mark G.
1993-01-01
A numerical analysis of the flowfield characteristics and the performance degradation of an airfoil with leading edge ice accretions was performed. The important fluid dynamic processes were identified and calculated. Among these were the leading edge separation bubble at low angles of attack, complete separation on the low pressure surface resulting in premature shell, drag rise due to the ice shape, and the effects of angle of attack on the separated flow field. Comparisons to experimental results were conducted to confirm these calculations. A computer code which solves the Navier-Stokes equations in two dimensions, ARC2D, was used to perform the calculations. A Modified Mixing Length turbulence model was developed to produce grids for several ice shape and airfoil combinations. Results indicate that the ability to predict overall performance characteristics, such as lift and drag, at low angles of attack is excellent. Transition location is important for accurately determining separation bubble shape. Details of the flowfield in and downstream of the separated regions requires some modifications. Calculations for the stalled airfoil indicate periodic shedding of vorticity that was generated aft of the ice accretion. Time averaged pressure values produce results which compare favorably with experimental information. A turbulence model which accounts for the history effects in the flow may be justified.
NASA Astrophysics Data System (ADS)
Mazaheri, K.; Nejati, A.; Chaharlang Kiani, K.; Taheri, R.
2016-07-01
A shock control bump (SCB) is a flow control method that uses local small deformations in a flexible wing surface to considerably reduce the strength of shock waves and the resulting wave drag in transonic flows. Most of the reported research is devoted to optimization in a single flow condition. Here, we have used a multi-point adjoint optimization scheme to optimize shape and location of the SCB. Practically, this introduces transonic airfoils equipped with the SCB that are simultaneously optimized for different off-design transonic flight conditions. Here, we use this optimization algorithm to enhance and optimize the performance of SCBs in two benchmark airfoils, i.e., RAE-2822 and NACA-64-A010, over a wide range of off-design Mach numbers. All results are compared with the usual single-point optimization. We use numerical simulation of the turbulent viscous flow and a gradient-based adjoint algorithm to find the optimum location and shape of the SCB. We show that the application of SCBs may increase the aerodynamic performance of an RAE-2822 airfoil by 21.9 and by 22.8 % for a NACA-64-A010 airfoil compared to the no-bump design in a particular flight condition. We have also investigated the simultaneous usage of two bumps for the upper and the lower surfaces of the airfoil. This has resulted in a 26.1 % improvement for the RAE-2822 compared to the clean airfoil in one flight condition.
A turbulence model for iced airfoils and its validation
NASA Technical Reports Server (NTRS)
Shin, Jaiwon; Chen, Hsun H.; Cebeci, Tuncer
1992-01-01
A turbulence model based on the extension of the algebraic eddy viscosity formulation of Cebeci and Smith developed for two dimensional flows over smooth and rough surfaces is described for iced airfoils and validated for computed ice shapes obtained for a range of total temperatures varying from 28 to -15 F. The validation is made with an interactive boundary layer method which uses a panel method to compute the inviscid flow and an inverse finite difference boundary layer method to compute the viscous flow. The interaction between inviscid and viscous flows is established by the use of the Hilbert integral. The calculated drag coefficients compare well with recent experimental data taken at the NASA-Lewis Icing Research Tunnel (IRT) and show that, in general, the drag increase due to ice accretion can be predicted well and efficiently.
Improved methods of vibration analysis of pretwisted, airfoil blades
NASA Technical Reports Server (NTRS)
Subrahmanyam, K. B.; Kaza, K. R. V.
1984-01-01
Vibration analysis of pretwisted blades of asymmetric airfoil cross section is performed by using two mixed variational approaches. Numerical results obtained from these two methods are compared to those obtained from an improved finite difference method and also to those given by the ordinary finite difference method. The relative merits, convergence properties and accuracies of all four methods are studied and discussed. The effects of asymmetry and pretwist on natural frequencies and mode shapes are investigated. The improved finite difference method is shown to be far superior to the conventional finite difference method in several respects. Close lower bound solutions are provided by the improved finite difference method for untwisted blades with a relatively coarse mesh while the mixed methods have not indicated any specific bound.
Theory and Experiment of Multielement Airfoils: A Comparison
NASA Technical Reports Server (NTRS)
Czerwiec, Ryan; Edwards, J. R.; Rumsey, C. L.; Hassan, H. A.
2000-01-01
A detailed comparison of computed and measured pressure distributions, velocity profiles, transition onset, and Reynolds shear stresses for multi-element airfoils is presented. It is shown that the transitional k-zeta model, which is implemented into CFL3D, does a good job of predicting pressure distributions, transition onset, and velocity profiles with the exception of velocities in the slat wake region. Considering the fact that the hot wire used was not fine enough to resolve Reynolds stresses in the boundary layer, comparisons of turbulence stresses varied from good to fair. It is suggested that the effects of unsteadiness be thoroughly evaluated before more complicated transition/turbulence models are used. Further, it is concluded that the present work presents a viable and economical method for calculating laminar/transitional/turbuient flows over complex shapes without user interface.
Exact solutions in oscillating airfoil theory
NASA Technical Reports Server (NTRS)
Williams, M. H.
1977-01-01
A result obtained by Williams (1977) for two-dimensional airfoils oscillating in an arbitrary subsonic parallel flowfield is reformulated to show that the pressure distribution induced by any deformation can be construed from the particular solutions for heaving and pitching motions. Specific formulas are presented for an oscillating control surface with a sealed gap.
Turbine airfoil with controlled area cooling arrangement
Liang, George
2010-04-27
A gas turbine airfoil (10) includes a serpentine cooling path (32) with a plurality of channels (34,42,44) fluidly interconnected by a plurality of turns (38,40) for cooling the airfoil wall material. A splitter component (50) is positioned within at least one of the channels to bifurcate the channel into a pressure-side channel (46) passing in between the outer wall (28) and the inner wall (30) of the pressure side (24) and a suction-side channel (48) passing in between the outer wall (28) and the inner wall (30) of the suction side (26) longitudinally downstream of an intermediate height (52). The cross-sectional area of the pressure-side channel (46) and suction-side channel (48) are thereby controlled in spite of an increasing cross-sectional area of the airfoil along its longitudinal length, ensuring a sufficiently high mach number to provide a desired degree of cooling throughout the entire length of the airfoil.
Near-wall serpentine cooled turbine airfoil
Lee, Ching-Pang
2014-10-28
A serpentine coolant flow path is formed by inner walls in a cavity between pressure and suction side walls of a turbine airfoil, the cavity partitioned by one or more transverse partitions into a plurality of continuous serpentine cooling flow streams each having a respective coolant inlet.
Parameterized post-Newtonian cosmology
NASA Astrophysics Data System (ADS)
Sanghai, Viraj A. A.; Clifton, Timothy
2017-03-01
Einstein’s theory of gravity has been extensively tested on solar system scales, and for isolated astrophysical systems, using the perturbative framework known as the parameterized post-Newtonian (PPN) formalism. This framework is designed for use in the weak-field and slow-motion limit of gravity, and can be used to constrain a large class of metric theories of gravity with data collected from the aforementioned systems. Given the potential of future surveys to probe cosmological scales to high precision, it is a topic of much contemporary interest to construct a similar framework to link Einstein’s theory of gravity and its alternatives to observations on cosmological scales. Our approach to this problem is to adapt and extend the existing PPN formalism for use in cosmology. We derive a set of equations that use the same parameters to consistently model both weak fields and cosmology. This allows us to parameterize a large class of modified theories of gravity and dark energy models on cosmological scales, using just four functions of time. These four functions can be directly linked to the background expansion of the universe, first-order cosmological perturbations, and the weak-field limit of the theory. They also reduce to the standard PPN parameters on solar system scales. We illustrate how dark energy models and scalar-tensor and vector-tensor theories of gravity fit into this framework, which we refer to as ‘parameterized post-Newtonian cosmology’ (PPNC).
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.
NASA Astrophysics Data System (ADS)
Gardner, A. D.; Klein, C.; Sachs, W. E.; Henne, U.; Mai, H.; Richter, K.
2014-09-01
Dynamic stall on a pitching OA209 airfoil in a wind tunnel is investigated at Mach 0.3 and 0.5 using high-speed pressure-sensitive paint (PSP) and pressure measurements. At Mach 0.3, the dynamic stall vortex was observed to propagate faster at the airfoil midline than at the wind-tunnel wall, resulting in a "bowed" vortex shape. At Mach 0.5, shock-induced stall was observed, with initial separation under the shock foot and subsequent expansion of the separated region upstream, downstream and along the breadth of the airfoil. No dynamic stall vortex could be observed at Mach 0.5. The investigation of flow control by blowing showed the potential advantages of PSP over pressure transducers for a complex three-dimensional flow.
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.
CLOUD PARAMETERIZATIONS, CLOUD PHYSICS, AND THEIR CONNECTIONS: AN OVERVIEW.
LIU,Y.; DAUM,P.H.; CHAI,S.K.; LIU,F.
2002-02-12
This paper consists of three parts. The first part is concerned with the parameterization of cloud microphysics in climate models. We demonstrate the crucial importance of spectral dispersion of the cloud droplet size distribution in determining radiative properties of clouds (e.g., effective radius), and underline the necessity of specifying spectral dispersion in the parameterization of cloud microphysics. It is argued that the inclusion of spectral dispersion makes the issue of cloud parameterization essentially equivalent to that of the droplet size distribution function, bringing cloud parameterization to the forefront of cloud physics. The second part is concerned with theoretical investigations into the spectral shape of droplet size distributions in cloud physics. After briefly reviewing the mainstream theories (including entrainment and mixing theories, and stochastic theories), we discuss their deficiencies and the need for a paradigm shift from reductionist approaches to systems approaches. A systems theory that has recently been formulated by utilizing ideas from statistical physics and information theory is discussed, along with the major results derived from it. It is shown that the systems formalism not only easily explains many puzzles that have been frustrating the mainstream theories, but also reveals such new phenomena as scale-dependence of cloud droplet size distributions. The third part is concerned with the potential applications of the systems theory to the specification of spectral dispersion in terms of predictable variables and scale-dependence under different fluctuating environments.
Unsteady Adjoint Approach for Design Optimization of Flapping Airfoils
NASA Technical Reports Server (NTRS)
Lee, Byung Joon; Liou, Meng-Sing
2012-01-01
This paper describes the work for optimizing the propulsive efficiency of flapping airfoils, i.e., improving the thrust under constraining aerodynamic work during the flapping flights by changing their shape and trajectory of motion with the unsteady discrete adjoint approach. For unsteady problems, it is essential to properly resolving time scales of motion under consideration and it must be compatible with the objective sought after. We include both the instantaneous and time-averaged (periodic) formulations in this study. For the design optimization with shape parameters or motion parameters, the time-averaged objective function is found to be more useful, while the instantaneous one is more suitable for flow control. The instantaneous objective function is operationally straightforward. On the other hand, the time-averaged objective function requires additional steps in the adjoint approach; the unsteady discrete adjoint equations for a periodic flow must be reformulated and the corresponding system of equations solved iteratively. We compare the design results from shape and trajectory optimizations and investigate the physical relevance of design variables to the flapping motion at on- and off-design conditions.
Turbine airfoil with laterally extending snubber having internal cooling system
Scribner, Carmen Andrew; Messmann, Stephen John; Marsh, Jan H.
2016-09-06
A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.
New airfoils for small horizontal axis wind turbines
Giguere, P.; Selig, M.S.
1997-12-31
In a continuing effort to enhance the performance of small energy systems, one root airfoil and three primary airfoils were specifically designed for small horizontal axis wind turbines. These airfoils are intended primarily for 1-10 kW variable-speed wind turbines for both conventional (tapered/twisted) or pultruded blades. The four airfoils were wind-tunnel tested at Reynolds numbers between 100,000 and 500,000. Tests with simulated leading-edge roughness were also conducted. The results indicate that small variable-speed wind turbines should benefit from the use of the new airfoils which provide enhanced lift-to-drag ratio performance as compared with previously existing airfoils.
Quiet airfoils for small and large wind turbines
Tangler, James L [Boulder, CO; Somers, Dan L [Port Matilda, PA
2012-06-12
Thick airfoil families with desirable aerodynamic performance with minimal airfoil induced noise. The airfoil families are suitable for a variety of wind turbine designs and are particularly well-suited for use with horizontal axis wind turbines (HAWTs) with constant or variable speed using pitch and/or stall control. In exemplary embodiments, a first family of three thick airfoils is provided for use with small wind turbines and second family of three thick airfoils is provided for use with very large machines, e.g., an airfoil defined for each of three blade radial stations or blade portions defined along the length of a blade. Each of the families is designed to provide a high maximum lift coefficient or high lift, to exhibit docile stalls, to be relatively insensitive to roughness, and to achieve a low profile drag.
Investigation of low-speed turbulent separated flow around airfoils
NASA Technical Reports Server (NTRS)
Wadcock, Alan J.
1987-01-01
Described is a low-speed wind tunnel experiment to measure the flowfield around a two-dimensional airfoil operating close to maximum lift. Boundary layer separation occurs on the upper surface at x/c=0.85. A three-component laser velocimeter, coupled with a computer-controlled data acquisition system, was used to obtain three orthogonal mean velocity components and three components of the Reynolds stress tensor in both the boundary layer and wake of the airfoil. Pressure distributions on the airfoil, skin friction distribution on the upper surface of the airfoil, and integral properties of the airfoil boudary layer are also documented. In addition to these near-field flow properties, static pressure distributions, both upstream and downstream from the airfoil and on the walls of the wind tunnel, are also presented.
Transonic airfoil analysis and design in nonuniform flow
NASA Technical Reports Server (NTRS)
Chang, J. F.; Lan, C. E.
1986-01-01
A nonuniform transonic airfoil code is developed for applications in analysis, inverse design and direct optimization involving an airfoil immersed in propfan slipstream. Problems concerning the numerical stability, convergence, divergence and solution oscillations are discussed. The code is validated by comparing with some known results in incompressible flow. A parametric investigation indicates that the airfoil lift-drag ratio can be increased by decreasing the thickness ratio. A better performance can be achieved if the airfoil is located below the slipstream center. Airfoil characteristics designed by the inverse method and a direct optimization are compared. The airfoil designed with the method of direct optimization exhibits better characteristics and achieves a gain of 22 percent in lift-drag ratio with a reduction of 4 percent in thickness.
NASA Technical Reports Server (NTRS)
Choo, Yung K.; Slater, John W.; Vickerman, Mary B.; VanZante, Judith F.; Wadel, Mary F. (Technical Monitor)
2002-01-01
Issues associated with analysis of 'icing effects' on airfoil and wing performances are discussed, along with accomplishments and efforts to overcome difficulties with ice. Because of infinite variations of ice shapes and their high degree of complexity, computational 'icing effects' studies using available software tools must address many difficulties in geometry acquisition and modeling, grid generation, and flow simulation. The value of each technology component needs to be weighed from the perspective of the entire analysis process, from geometry to flow simulation. Even though CFD codes are yet to be validated for flows over iced airfoils and wings, numerical simulation, when considered together with wind tunnel tests, can provide valuable insights into 'icing effects' and advance our understanding of the relationship between ice characteristics and their effects on performance degradation.
Application of numerical optimization to the design of supercritical airfoils without drag-creep
NASA Technical Reports Server (NTRS)
Hicks, R. M.; Vanderplaats, G. N.
1977-01-01
Recent applications of numerical optimization to the design of advanced airfoils for transonic aircraft have shown that low-drag sections can be developed for a given design Mach number without an accompanying drag increase at lower Mach numbers. This is achieved by imposing a constraint on the drag coefficient at an off-design Mach number while the drag at the design Mach number is the objective function. Such a procedure doubles the computation time over that for single design-point problems, but the final result is worth the increased cost of computation. The ability to treat such multiple design-point problems by numerical optimization has been enhanced by the development of improved airfoil shape functions. Such functions permit a considerable increase in the range of profiles attainable during the optimization process.
Separated transonic airfoil flow calculations with a nonequilibrium turbulence model
NASA Technical Reports Server (NTRS)
King, L. S.; Johnson, D. A.
1985-01-01
Navier-Stokes transonic airfoil calculations based on a recently developed nonequilibrium, turbulence closure model are presented for a supercritical airfoil section at transonic cruise conditions and for a conventional airfoil section at shock-induced stall conditions. Comparisons with experimental data are presented which show that this nonequilibrium closure model performs significantly better than the popular Baldwin-Lomax and Cebeci-Smith equilibrium algebraic models when there is boundary-layer separation that results from the inviscid-viscous interactions.
S825 and S826 Airfoils: 1994--1995
Somers, D. M.
2005-01-01
A family of airfoils, the S825 and S826, for 20- to 40-meter, variable-speed and variable-pitch (toward feather), horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moments and the airfoil thicknesses have been satisfied. The airfoils should exhibit docile stalls.
NASA Technical Reports Server (NTRS)
Swanson, Robert S.; Schuldenfrei, Marvin J.
1940-01-01
An investigation has been made in the NACA 7- by 10-foot wind tunnel of a large chord NACA 27-212 airfoil with a 20% chord split flap and with two arrangements of a 25.66% chord slotted flap to determine the section lift characteristics as affected by flap deflection for the split flap and as affected by flap deflection, flap position, and slot shape for the slotted flap. For the two arrangements of the slotted flap, the flap positions for maximum section lift are given. Comparable data on the NACA 23012 airfoil equipped with similar flaps are also given. On the basis of maximum section lift coefficient, the slotted flap with an easy slot entry was slightly better than either the split flap or the slotted flap with a sharp slot entry. With both types of flap the decrease in the angle of attack, for maximum section lift coefficient, with flap deflection is large for the NACA 27-212 airfoil as compared with the NACA 23012 airfoil. Also with both flaps, the maximum section lift coefficient obtained with flaps is much lower for the NACA 27-212 airfoil than for the NACA 23012 airfoil.
Turbine airfoil with ambient cooling system
Campbell, Jr, Christian X.; Marra, John J.; Marsh, Jan H.
2016-06-07
A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.
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.
Turbine engine airfoil and platform assembly
Campbell, Christian X [Oviedo, FL; James, Allister W [Chuluota, FL; Morrison, Jay A [Oviedo, FL
2012-07-31
A turbine airfoil (22A) is formed by a first process using a first material. A platform (30A) is formed by a second process using a second material that may be different from the first material. The platform (30A) is assembled around a shank (23A) of the airfoil. One or more pins (36A) extend from the platform into holes (28) in the shank (23A). The platform may be formed in two portions (32A, 34A) and placed around the shank, enclosing it. The two platform portions may be bonded to each other. Alternately, the platform (30B) may be cast around the shank (23B) using a metal alloy with better castability than that of the blade and shank, which may be specialized for thermal tolerance. The pins (36A-36D) or holes for them do not extend to an outer surface (31) of the platform, avoiding stress concentrations.
Turbomachinery Airfoil Design Optimization Using Differential Evolution
NASA Technical Reports Server (NTRS)
Madavan, Nateri K.; Biegel, Bryan (Technical Monitor)
2002-01-01
An aerodynamic design optimization procedure that is based on a evolutionary algorithm known at Differential Evolution is described. Differential Evolution is a simple, fast, and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems, including highly nonlinear systems with discontinuities and multiple local optima. The method is combined with a Navier-Stokes solver that evaluates the various intermediate designs and provides inputs to the optimization procedure. An efficient constraint handling mechanism is also incorporated. Results are presented for the inverse design of a turbine airfoil from a modern jet engine and compared to earlier methods. The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated. Substantial reductions in the overall computing time requirements are achieved by using the algorithm in conjunction with neural networks.
Turbomachinery Airfoil Design Optimization Using Differential Evolution
NASA Technical Reports Server (NTRS)
Madavan, Nateri K.; Biegel, Bryan A. (Technical Monitor)
2002-01-01
An aerodynamic design optimization procedure that is based on a evolutionary algorithm known at Differential Evolution is described. Differential Evolution is a simple, fast, and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems, including highly nonlinear systems with discontinuities and multiple local optima. The method is combined with a Navier-Stokes solver that evaluates the various intermediate designs and provides inputs to the optimization procedure. An efficient constraint handling mechanism is also incorporated. Results are presented for the inverse design of a turbine airfoil from a modern jet engine. The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated. Substantial reductions in the overall computing time requirements are achieved by using the algorithm in conjunction with neural networks.
Low Reynolds number airfoil survey, volume 1
NASA Technical Reports Server (NTRS)
Carmichael, B. H.
1981-01-01
The differences in flow behavior two dimensional airfoils in the critical chordlength Reynolds number compared with lower and higher Reynolds number are discussed. The large laminar separation bubble is discussed in view of its important influence on critical Reynolds number airfoil behavior. The shortcomings of application of theoretical boundary layer computations which are successful at higher Reynolds numbers to the critical regime are discussed. The large variation in experimental aerodynamic characteristic measurement due to small changes in ambient turbulence, vibration, and sound level is illustrated. The difficulties in obtaining accurate detailed measurements in free flight and dramatic performance improvements at critical Reynolds number, achieved with various types of boundary layer tripping devices are discussed.
Aerodynamic properties of thick airfoils II
NASA Technical Reports Server (NTRS)
Norton, F H; Bacon, D L
1923-01-01
This investigation is an extension of NACA report no. 75 for the purpose of studying the effect of various modifications in a given wing section, including changes in thickness, height of lower camber, taper in thickness, and taper in plan form with special reference to the development of thick, efficient airfoils. The method consisted in testing the wings in the NACA 5-foot wind tunnel at speeds up to 50 meters (164 feet) per second while they were being supported on a new type of wire balance. Some of the airfoils developed showed results of great promise. For example, one wing (no. 81) with a thickness in the center of 4.5 times that of the U. S. A. 16 showed both uniformly high efficiency and a higher maximum lift than this excellent section. These thick sections will be especially useful on airplanes with cantilever construction. (author)
Experiments on airfoils with trailing edge cut away
NASA Technical Reports Server (NTRS)
Ackeret, J
1927-01-01
Airfoils with their trailing edge cut away are often found on aircraft, as the fins on the hulls of flying boats and the central section of the wings for affording better visibility. It was therefore of some interest to discover the effect of such cutaways on the lift and drag and on the position of the center of pressure. For this purpose, systematic experiments were performed on two different airfoils, a symmetrical airfoil and an airfoil of medium thickness, with successive shortenings of their chords.
Damping element for reducing the vibration of an airfoil
Campbell, Christian X; Marra, John J
2013-11-12
An airfoil (10) is provided with a tip (12) having an opening (14) to a center channel (24). A damping element (16) is inserted within the opening of the center channel, to reduce an induced vibration of the airfoil. The mass of the damping element, a spring constant of the damping element within the center channel, and/or a mounting location (58) of the damping element within the center channel may be adjustably varied, to shift a resonance frequency of the airfoil outside a natural operating frequency of the airfoil.
Transonic airfoil analysis and design using Cartesian coordinates
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1975-01-01
An inverse numerical technique for designing transonic airfoils having a prescribed pressure distribution is presented. The method uses the full potential equation, inverse boundary conditions, and Cartesian coordinates. It includes simultaneous airfoil update and utilizes a direct-inverse approach that permits a logical method for controlling trailing edge closure. The method can also be used for the analysis of flowfields about specified airfoils. Comparison with previous results shows that accurate results can be obtained with a Cartesian grid. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.
Numerical investigation of acoustic radiation from vortex-airfoil interaction
NASA Astrophysics Data System (ADS)
Legault, Anne; Ji, Minsuk; Wang, Meng
2012-11-01
Numerical simulations of vortices interacting with a NACA 0012 airfoil and a flat-plate airfoil at zero angle of attack are carried out to assess the applicability and accuracy of classical theories. Unsteady lift and sound are computed and compared with the predictions by theories of Sears and Amiet, which assume a thin-plate airfoil in an inviscid flow. A Navier-Stokes solver is used in the simulations, and therefore viscous effects are taken into consideration. For the thin-plate airfoil, the effect of viscosity is negligible. For a NACA 0012 airfoil, the viscous contribution to the unsteady lift and sound mainly comes from coherent vortex shedding in the wake of the airfoil and the interaction of the incoming vortices with the airfoil wake, which become stronger at higher Reynolds numbers for a 2-D laminar flow. When the flow is turbulent at chord Reynolds number of 4 . 8 ×105 , however, the viscous contribution becomes negligible as coherent vortex shedding is not present. Sound radiation from vortex-airfoil interaction at turbulent Reynolds numbers is computed numerically via Lighthill's theory and the result is compared with the predictions of Amiet and Curle. The effect of the airfoil thickness is also examined. Supported by ONR Grant N00014-09-1-1088.
Comparisons of Theoretical Methods for Predicting Airfoil Aerodynamic Characteristics
2010-08-01
Airfoil ,” Airfoils , U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D-107, August 2010. [2] Somers, D.M. and...Maughmer, M.D., “Design and Experimental Results for the S407 Airfoil ,” U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D...S414 Airfoil ,” U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D-112, August 2010. [5] Somers, D.M. and Maughmer
Investigation of the Boundary Layer Behavior on Turbine Airfoils.
1979-08-01
turbine airfoil cascade . The airfoil profile was based on a turbine blade design used by Lander ’’4 and employed in previous wake studies by Cox and...simulate the wake from upstream turning vanes or blades , a circular cylinder was placed upstream of the centra l or test airfoil . The displacement of this...of turbine airfoil cascade model s by Cox and Han 15 are very much evident in the graph . It might be noted that the blade stag- nation points are at
Development and testing of airfoils for high-altitude aircraft
NASA Technical Reports Server (NTRS)
Drela, Mark (Principal Investigator)
1996-01-01
Specific tasks included airfoil design; study of airfoil constraints on pullout maneuver; selection of tail airfoils; examination of wing twist; test section instrumentation and layout; and integrated airfoil/heat-exchanger tests. In the course of designing the airfoil, specifically for the APEX test vehicle, extensive studies were made over the Mach and Reynolds number ranges of interest. It is intended to be representative of airfoils required for lightweight aircraft operating at extreme altitudes, which is the primary research objective of the APEX program. Also considered were thickness, pitching moment, and off-design behavior. The maximum ceiling parameter M(exp 2)C(sub L) value achievable by the Apex-16 airfoil was found to be a strong constraint on the pullout maneuver. The NACA 1410 and 2410 airfoils (inverted) were identified as good candidates for the tail, with predictable behavior at low Reynolds numbers and good tolerance to flap deflections. With regards to wing twist, it was decided that a simple flat wing was a reasonable compromise. The test section instrumentation consisted of surface pressure taps, wake rakes, surface-mounted microphones, and skin-friction gauges. Also, a modest wind tunnel test was performed for an integrated airfoil/heat-exchanger configuration, which is currently on Aurora's 'Theseus' aircraft. Although not directly related to the APEX tests, the aerodynamics or heat exchangers has been identified as a crucial aspect of designing high-altitude aircraft and hence is relevant to the ERAST program.
NASA Technical Reports Server (NTRS)
Mutterperl, William
1944-01-01
A method of conformal transformation is developed that maps an airfoil into a straight line, the line being chosen as the extended chord line of the airfoil. The mapping is accomplished by operating directly with the airfoil ordinates. The absence of any preliminary transformation is found to shorten the work substantially over that of previous methods. Use is made of the superposition of solutions to obtain a rigorous counterpart of the approximate methods of thin-airfoils theory. The method is applied to the solution of the direct and inverse problems for arbitrary airfoils and pressure distributions. Numerical examples are given. Applications to more general types of regions, in particular to biplanes and to cascades of airfoils, are indicated. (author)
Streamwise Oscillation of Airfoils into Reverse Flow
NASA Astrophysics Data System (ADS)
Granlund, Kenneth; Jones, Anya; Ol, Michael
2015-11-01
An airfoil in freestream is oscillated in streamwise direction to cyclically enter reverse flow. Measured lift is compared to analytical blade element theories. Advance ratio, reduced frequency and angle of attack is varied within those typical for helicopters. Experimental results reveal that lift does not become negative in the flow reversal part, contradicting one theory and supported by another. Flow visualization reveal the leading edge vortex advecting against the freestream to a point in front of the leading edge.
Linearized propulsion theory of flapping airfoils revisited
NASA Astrophysics Data System (ADS)
Fernandez-Feria, R.
2016-12-01
A vortical impulse theory is used to compute the thrust force of a plunging and pitching airfoil in forward flight at high Reynolds numbers within the framework of linear potential flow theory. The result is significantly different from the classical one of Garrick, which considered only two effects, the leading-edge suction and the projection in the flight direction of the pressure force on the airfoil. By taking into account the complete vorticity distribution on the airfoil and the wake the mean thrust coefficient contains, in addition to the pressure force projection term, a new term that generalizes the leading-edge suction term in Garrick's theory. This term depends on Theodorsen function C (k ) and on a new complex function C1(k ) of the reduced frequency k . The main qualitative difference with Garrick's theory is that the propulsive efficiency, or ratio of the mean thrust power and the mean input power required to drive the airfoil, tends to zero as the reduced frequency increases to infinity (as k-1), in contrast to Garrick's propulsive efficiency that tends to a constant (1 /2 ). Consequently, for pure pitching and combined pitching and plunging motions, the maximum of the propulsive efficiency is not reached as k →∞ like in Garrick's theory, but at a finite value of the reduced frequency that depends on the remaining nondimensional parameters. The present analytical results are in good agreement, for small amplitude oscillations, with numerical results from unsteady panel methods, and with experimental data and numerical results from the Navier-Stokes equations, except for small reduced frequencies where viscous effects are obviously important.
Turbulent Flow over Rough Turbine Airfoils.
1985-08-01
SUBJECT TERMS (Continue on reverse if necessary and identify by block number) FIELD GROUP SUB. GR. Turbine blades ’ vanes ; surface roughness...turbulent boundary layer over rough turbine vanes or blades is developed. A new formulation of the mixing length model, expressed in the velocity-space...A-163 005 TURBULENT FLOW OVER ROUGH TURBINE AIRFOILS (U) OHIO 1/ STATE UNIV RESEARCH FOUNDATION COLUMBUS L S HAN AUG B5 OSURF-76357/?i4467 AFWL-TR-95
Tail Rotor Airfoils Stabilize Helicopters, Reduce Noise
NASA Technical Reports Server (NTRS)
2010-01-01
Founded by former Ames Research Center engineer Jim Van Horn, Van Horn Aviation of Tempe, Arizona, built upon a Langley Research Center airfoil design to create a high performance aftermarket tail rotor for the popular Bell 206 helicopter. The highly durable rotor has a lifetime twice that of the original equipment manufacturer blade, reduces noise by 40 percent, and displays enhanced performance at high altitudes. These improvements benefit helicopter performance for law enforcement, military training, wildfire and pipeline patrols, and emergency medical services.
NASA Technical Reports Server (NTRS)
Taylor, Arthur C., III; Newman, James C., III; Barnwell, Richard W.
1997-01-01
A three-dimensional unstructured grid approach to aerodynamic shape sensitivity analysis and design optimization has been developed and is extended to model geometrically complex configurations. The advantage of unstructured grids (when compared with a structured-grid approach) is their inherent ability to discretize irregularly shaped domains with greater efficiency and less effort. Hence, this approach is ideally suited for geometrically complex configurations of practical interest. In this work the nonlinear Euler equations are solved using an upwind, cell-centered, finite-volume scheme. The discrete, linearized systems which result from this scheme are solved iteratively by a preconditioned conjugate-gradient-like algorithm known as GMRES for the two-dimensional geometry and a Gauss-Seidel algorithm for the three-dimensional; similar procedures are used to solve the accompanying linear aerodynamic sensitivity equations in incremental iterative form. As shown, this particular form of the sensitivity equation makes large-scale gradient-based aerodynamic optimization possible by taking advantage of memory efficient methods to construct exact Jacobian matrix-vector products. Simple parameterization techniques are utilized for demonstrative purposes. Once the surface has been deformed, the unstructured grid is adapted by considering the mesh as a system of interconnected springs. Grid sensitivities are obtained by differentiating the surface parameterization and the grid adaptation algorithms with ADIFOR (which is an advanced automatic-differentiation software tool). To demonstrate the ability of this procedure to analyze and design complex configurations of practical interest, the sensitivity analysis and shape optimization has been performed for a two-dimensional high-lift multielement airfoil and for a three-dimensional Boeing 747-200 aircraft.
Synthesis of Entrainment and Detrainment formulations for Convection Parameterizations
NASA Astrophysics Data System (ADS)
Siebesma, P.
2015-12-01
Mixing between convective clouds and its environment, usually parameterized in terms of entrainment and detrainment, are among the most important processes that determine the strength of the climate model sensitivity. This notion has led to a renaissance of research in exploring the mechanisms of these mixing processes and, as a result, to a wide range of seemingly different parameterized formulations. In this study we are aiming to synthesize these results as to offer a solid framework for use in parameterized formulations of convection. Detailed LES analyses in which clouds are subsampled according to their size show that entrainment rates are inversely proportional to the typical cloud radius, in accordance with original entraining plume models. These results can be shown analytically to be consistent with entrainment rate formulations of cloud ensembles that decrease inversely proportional with height, by making only mild assumptions on the shape of the associated cloud size distribution. In addition there are additional dependencies of the entrainment rates on the environmental thermodynamics such as the relative humidity and stability but these are of second order. In contrast detrainment rates do depend to first order on the environmental thermodynamics such as relative humidity and stability. This can be understood by realizing that i) the details of the cloud size distribution do depend on these environmental factors and ii) that detrainment rates have a much stronger dependency on the shape of the cloud size distribution than entrainment rates.
Infrared radiation parameterizations in numerical climate models
NASA Technical Reports Server (NTRS)
Chou, Ming-Dah; Kratz, David P.; Ridgway, William
1991-01-01
This study presents various approaches to parameterizing the broadband transmission functions for utilization in numerical climate models. One-parameter scaling is applied to approximate a nonhomogeneous path with an equivalent homogeneous path, and the diffuse transmittances are either interpolated from precomputed tables or fit by analytical functions. Two-parameter scaling is applied to parameterizing the carbon dioxide and ozone transmission functions in both the lower and middle atmosphere. Parameterizations are given for the nitrous oxide and methane diffuse transmission functions.
Pressure Distribution Over Airfoils at High Speeds
NASA Technical Reports Server (NTRS)
Briggs, L J; Dryden, H L
1927-01-01
This report deals with the pressure distribution over airfoils at high speeds, and describes an extension of an investigation of the aerodynamic characteristics of certain airfoils which was presented in NACA Technical Report no. 207. The results presented in report no. 207 have been confirmed and extended to higher speeds through a more extensive and systematic series of tests. Observations were also made of the air flow near the surface of the airfoils, and the large changes in lift coefficients were shown to be associated with a sudden breaking away of the flow from the upper surface. The tests were made on models of 1-inch chord and comparison with the earlier measurements on models of 3-inch chord shows that the sudden change in the lift coefficient is due to compressibility and not to a change in the Reynolds number. The Reynolds number still has a large effect, however, on the drag coefficient. The pressure distribution observations furnish the propeller designer with data on the load distribution at high speeds, and also give a better picture of the air-flow changes.
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.
Wake structure of a deformable Joukowski airfoil
NASA Astrophysics Data System (ADS)
Ysasi, Adam; Kanso, Eva; Newton, Paul K.
2011-10-01
We examine the vortical wake structure shed from a deformable Joukowski airfoil in an unbounded volume of inviscid and incompressible fluid. The deformable airfoil is considered to model a flapping fish. The vortex shedding is accounted for using an unsteady point vortex model commonly referred to as the Brown-Michael model. The airfoil’s deformations and rotations are prescribed in terms of a Jacobi elliptic function which exhibits, depending on a dimensionless parameter m, a range of periodic behaviors from sinusoidal to a more impulsive type flapping. Depending on the parameter m and the Strouhal number, one can identify five distinct wake structures, ranging from arrays of isolated point vortices to vortex dipoles and tripoles shed into the wake with every half-cycle of the airfoil flapping motion. We describe these regimes in the context of other published works which categorize wake topologies, and speculate on the importance of these wake structures in terms of periodic swimming and transient maneuvers of fish.
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.
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
Parameterization of solar flare dose
Lamarche, A.H.; Poston, J.W.
1996-12-31
A critical aspect of missions to the moon or Mars will be the safety and health of the crew. Radiation in space is a hazard for astronauts, especially high-energy radiation following certain types of solar flares. A solar flare event can be very dangerous if astronauts are not adequately shielded because flares can deliver a very high dose in a short period of time. The goal of this research was to parameterize solar flare dose as a function of time to see if it was possible to predict solar flare occurrence, thus providing a warning time. This would allow astronauts to take corrective action and avoid receiving a dose greater than the recommended limit set by the National Council on Radiation Protection and Measurements (NCRP).
A smooth chaotic map with parameterized shape and symmetry
NASA Astrophysics Data System (ADS)
Chaves, Daniel P. B.; Souza, Carlos E. C.; Pimentel, Cecilio
2016-12-01
We introduce in this paper a new chaotic map with dynamical properties controlled by two free parameters. The map definition is based on the hyperbolic tangent function, so it is called the tanh map. We demonstrate that the Lyapunov exponent of the tanh map is robust, remaining practically unaltered with the variation of its parameters. As the main application, we consider a chaotic communication system based on symbolic dynamics with advantages over current approaches that use piecewise linear maps. In this context, we propose a new measure, namely, the spread rate, to study the local structure of the chaotic dynamics of a one-dimensional chaotic map.
Visibility Parameterization For Forecasting Model Applications
NASA Astrophysics Data System (ADS)
Gultepe, I.; Milbrandt, J.; Binbin, Z.
2010-07-01
In this study, the visibility parameterizations developed during Fog Remote Sensing And Modeling (FRAM) projects, conducted in central and eastern Canada, will be summarized and their use for forecasting/nowcasting applications will be discussed. Parameterizations developed for reductions in visibility due to 1) fog, 2) rain, 3) snow, and 4) relative humidity (RH) during FRAM will be given and uncertainties in the parameterizations will be discussed. Comparisons made between Canadian GEM NWP model (with 1 and 2.5 km horizontal grid spacing) and observations collected during the Science of Nowcasting Winter Weather for Vancouver 2010 (SNOW-V10) project and FRAM projects, using the new parameterizations, will be given Observations used in this study were obtained using a fog measuring device (FMD) for fog parameterization, a Vaisala all weather precipitation sensor called FD12P for rain and snow parameterizations and visibility measurements, and a total precipitation sensor (TPS), and distrometers called OTT ParSiVel and Laser Precipitation Measurement (LPM) for rain/snow particle spectra. The results from the three SNOW-V10 sites suggested that visibility values given by the GEM model using the new parameterizations were comparable with observed visibility values when model based input parameters such as liquid water content, RH, and precipitation rate for visibility parameterizations were predicted accurately.
A semi-empirical airfoil stall noise model based on surface pressure measurements
NASA Astrophysics Data System (ADS)
Bertagnolio, Franck; Madsen, Helge Aa.; Fischer, Andreas; Bak, Christian
2017-01-01
This work is concerned with the experimental study of airfoil stall and the modelling of stall noise. Using pressure taps and high-frequency surface pressure microphones flush-mounted on airfoils measured in wind tunnels and on an operating wind turbine blade, the characteristics of stall are analyzed. This study shows that the main quantities of interest, namely convection velocity, spatial correlation and surface pressure spectra, can be scaled highlighting the universal nature of stall independently of airfoil shapes and flow conditions, although within a certain range of experimental conditions. Two main regimes for the scaling of the correlation lengths and the surface pressure spectra, depending on the Reynolds number of the flow, can be distinguished. These results are used to develop a model for the surface pressure spectra within the detached flow region valid for Reynolds numbers ranging from 1 ×106 to 6 ×106. Subsequently, this model is used to derive a model for stall noise. Modelled noise spectra are compared with experimental data measured in anechoic wind tunnels with reasonably satisfactory agreement.
Vapor deposition on doublet airfoil substrates: Control of coating thickness and microstructure
Rodgers, Theron M.; Zhao, Hengbei; Wadley, Haydn N. G.
2015-11-15
Gas jet assisted vapor deposition processes for depositing coatings are conducted at higher pressures than conventional physical vapor deposition methods, and have shown promise for coating complex shaped substrates including those with non-line-of-sight (NLS) regions on their surface. These regions typically receive vapor atoms at a lower rate and with a wider incident angular distribution than substrate regions in line-of-sight (LS) of the vapor source. To investigate the coating of such substrates, the thickness and microstructure variation along the inner (curved) surfaces of a model doublet airfoil containing both LS and NLS regions has been investigated. Results from atomistic simulations and experiments confirm that the coating's thickness is thinner in flux-shadowed regions than in other regions for all the coating processes investigated. They also indicated that the coatings columnar microstructure and pore volume fraction vary with surface location through the LS to NLS transition zone. A substrate rotation strategy for optimizing the thickness over the entire doublet airfoil surface was investigated, and led to the identification of a process that resulted in only small variation of coating thickness, columnar growth angle, and pore volume fraction on all doublet airfoil surfaces.
Local and Total Density Measurements in Ice Shapes
NASA Technical Reports Server (NTRS)
Vargas, Mario; Broughton, Howard; Sims, James J.; Bleeze, Brian; Gaines, Vatanna
2005-01-01
Preliminary measurements of local and total densities inside ice shapes were obtained from ice shapes grown in the NASA Glenn Research Tunnel for a range of glaze ice, rime ice, and mixed phase ice conditions on a NACA 0012 airfoil at 0 angle of attack. The ice shapes were removed from the airfoil and a slice of ice 3 mm thick was obtained using a microtome. The resulting samples were then x-rayed to obtain a micro-radiography, the film was digitized, and image processing techniques were used to extract the local and total density values.
Sealing apparatus for airfoils of gas turbine engines
Jones, R.B.
1998-05-19
An improved airfoil tip sealing apparatus is disclosed wherein brush seals are attached to airfoil tips with the distal ends of the brush seal fibers sealingly contacting opposing wall surfaces. Embodiments for variable vanes, stators and both cooled and uncooled turbine blade applications are disclosed. 17 figs.
Sealing apparatus for airfoils of gas turbine engines
Jones, Russell B.
1998-01-01
An improved airfoil tip sealing apparatus is disclosed wherein brush seals are attached to airfoil tips with the distal ends of the brush seal fibers sealingly contacting opposing wall surfaces. Embodiments for variable vanes, stators and both cooled and uncooled turbine blade applications are disclosed.
Analytical studies of new airfoils for wind turbines
NASA Technical Reports Server (NTRS)
Wentz, W. H., Jr.; Calhoun, J. T.
1981-01-01
Computer studies were conducted to analyze the potential gains associated with utilizing new airfoils for large wind turbine rotor blades. Attempts to include 3-dimensional stalling effects were inconclusive. It is recommended that blade pressure measurements be made to clarify the nature of blade stalling. It is also recommended that new laminar flow airfoils be used as rotor blade sections.
Design procedure for low-drag subsonic airfoils
NASA Technical Reports Server (NTRS)
Peterson, J. B.; Chen, A. B.
1975-01-01
Airfoil has least amount of drag under given restrictions of boundary layer transition position, lift coefficient, thickness ratio, and Reynolds number based on airfoil chord. It is suitable for use as wing and propeller aircraft sections operating at subsonic speeds and for hydrofoil sections and blades for fans, compressors, turbines, and windmills.
TRANDES: A FORTRAN program for transonic airfoil analysis or design
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1977-01-01
A program called TRANDES is presented that is used for the analysis of steady, irrotational transonic flow over specified two-dimensional airfoils in free air or for the design of airfoils having a prescribed pressure distribution, including the effects of weak viscous interaction. Instructions on program usage, listings of the program, and sample cases are given.
Development of drive mechanism for an oscillating airfoil
NASA Technical Reports Server (NTRS)
Sticht, Clifford D.
1988-01-01
The design and development of an in-draft wind tunnel test section which will be used to study the dynamic stall of airfoils oscillating in pitch is described. The hardware developed comprises a spanned airfoil between schleiren windows, a four bar linkage, flywheels, a drive system and a test section structure.
An approach for parameterizing mesoscale precipitating systems
Weissbluth, M.J.; Cotton, W.R.
1991-12-31
A cumulus parameterization laboratory has been described which uses a reference numerical model to fabricate, calibrate and verify a cumulus parameterization scheme suitable for use in mesoscale models. Key features of this scheme include resolution independence and the ability to provide hydrometeor source functions to the host model. Thus far, only convective scale drafts have been parameterized, limiting the use of the scheme to those models which can resolve the mesoscale circulations. As it stands, the scheme could probably be incorporated into models having a grid resolution greater than 50 km with results comparable to the existing schemes for the large-scale models. We propose, however, to quantify the mesoscale circulations through the use of the cumulus parameterization laboratory. The inclusion of these mesoscale drafts in the existing scheme will hopefully allow the correct parameterization of the organized mesoscale precipitating systems.
An approach for parameterizing mesoscale precipitating systems
Weissbluth, M.J.; Cotton, W.R.
1991-01-01
A cumulus parameterization laboratory has been described which uses a reference numerical model to fabricate, calibrate and verify a cumulus parameterization scheme suitable for use in mesoscale models. Key features of this scheme include resolution independence and the ability to provide hydrometeor source functions to the host model. Thus far, only convective scale drafts have been parameterized, limiting the use of the scheme to those models which can resolve the mesoscale circulations. As it stands, the scheme could probably be incorporated into models having a grid resolution greater than 50 km with results comparable to the existing schemes for the large-scale models. We propose, however, to quantify the mesoscale circulations through the use of the cumulus parameterization laboratory. The inclusion of these mesoscale drafts in the existing scheme will hopefully allow the correct parameterization of the organized mesoscale precipitating systems.
Airfoil family design for large offshore wind turbine blades
NASA Astrophysics Data System (ADS)
Méndez, B.; Munduate, X.; San Miguel, U.
2014-06-01
Wind turbine blades size has scaled-up during last years due to wind turbine platform increase especially for offshore applications. The EOLIA project 2007-2010 (Spanish Goverment funded project) was focused on the design of large offshore wind turbines for deep waters. The project was managed by ACCIONA Energia and the wind turbine technology was designed by ACCIONA Windpower. The project included the design of a wind turbine airfoil family especially conceived for large offshore wind turbine blades, in the order of 5MW machine. Large offshore wind turbines suffer high extreme loads due to their size, in addition the lack of noise restrictions allow higher tip speeds. Consequently, the airfoils presented in this work are designed for high Reynolds numbers with the main goal of reducing blade loads and mantainig power production. The new airfoil family was designed in collaboration with CENER (Spanish National Renewable Energy Centre). The airfoil family was designed using a evolutionary algorithm based optimization tool with different objectives, both aerodynamic and structural, coupled with an airfoil geometry generation tool. Force coefficients of the designed airfoil were obtained using the panel code XFOIL in which the boundary layer/inviscid flow coupling is ineracted via surface transpiration model. The desing methodology includes a novel technique to define the objective functions based on normalizing the functions using weight parameters created from data of airfoils used as reference. Four airfoils have been designed, here three of them will be presented, with relative thickness of 18%, 21%, 25%, which have been verified with the in-house CFD code, Wind Multi Block WMB, and later validated with wind tunnel experiments. Some of the objectives for the designed airfoils concern the aerodynamic behavior (high efficiency and lift, high tangential coefficient, insensitivity to rough conditions, etc.), others concern the geometry (good for structural design
Multiple element airfoils optimized for maximum lift coefficient.
NASA Technical Reports Server (NTRS)
Ormsbee, A. I.; Chen, A. W.
1972-01-01
Optimum airfoils in the sense of maximum lift coefficient are obtained for incompressible fluid flow at large Reynolds number. The maximum lift coefficient is achieved by requiring that the turbulent skin friction be zero in the pressure rise region on the airfoil upper surface. Under this constraint, the pressure distribution is optimized. The optimum pressure distribution is a function of Reynolds number and the trailing edge velocity. Geometries of those airfoils which will generate these optimum pressure distributions are obtained using a direct-iterative method which is developed in this study. This method can be used to design airfoils consisting of any number of elements. Numerical examples of one- and two-element airfoils are given. The maximum lift coefficients obtained range from 2 to 2.5.
Transition effects on compressible dynamic stall of transiently pitching airfoils
NASA Technical Reports Server (NTRS)
Wilder, M. C.; Chandrasekhara, M. S.; Carr, L. W.
1993-01-01
Experimental results and analysis of the effects of boundary layer tripping on dynamic stall of a transiently pitching airfoil are presented. At low Mach numbers, the tripped airfoil exhibits qualitative similarity with the behavior of the untripped airfoil. However, the local supersonic flow at Mach numbers greater than 0.3 is significantly modified by the trip leading to vastly different shock/boundary layer interactions, dynamic stall onset and vortex formation angles. The formation of the laminar separation bubble is found to have a favorable influence in delaying dynamic stall on the untripped airfoil flow. In both Mach number regimes, the tripped flow actually stalls at slightly lower angles of attack. Further experimentation with three trips on an oscillating airfoil showed that the dynamic stall process is very sensitive to the state of the turbulence in the boundary layer. This sensitivity points to a need for finer turbulence modeling techniques for use in dynamic stall computations.
Wind tunnel test of the S814 thick root airfoil
Somers, D.M.; Tangler, J.L.
1996-11-01
The objective of this wind-tunnel test was to verify the predictions of the Eppler Airfoil Design and Analysis Code for a very thick airfoil having a high maximum lift coefficient designed to be largely insensitive to leading-edge roughness effects. The 24 percent thick S814 airfoil was designed with these characteristics to accommodate aerodynamic and structural considerations for the root region of a wind-turbine blade. In addition, the airfoil`s maximum lift-to-drag ratio was designed to occur at a high lift coefficient. To accomplish the objective, a two-dimensional wind tunnel test of the S814 thick root airfoil was conducted in January 1994 in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. Data were obtained with transition free and transition fixed for Reynolds numbers of 0.7, 1.0, 1.5, 2.0, and 3.0 {times} 10{sup 6}. For the design Reynolds number of 1.5 {times} 10{sup 6}, the maximum lift coefficient with transition free is 1.32, which satisfies the design specification. However, this value is significantly lower than the predicted maximum lift coefficient of almost 1.6. With transition fixed at the leading edge, the maximum lift coefficient is 1.22. The small difference in maximum lift coefficient between the transition-free and transition-fixed conditions demonstrates the airfoil`s minimal sensitivity to roughness effects. The S814 root airfoil was designed to complement existing NREL low maximum-lift-coefficient tip-region airfoils for rotor blades 10 to 15 meters in length.
Reversible airfoils for stopped rotors in high speed flight
NASA Astrophysics Data System (ADS)
Niemiec, Robert; Jacobellis, George; Gandhi, Farhan
2014-10-01
This study starts with the design of a reversible airfoil rib for stopped-rotor applications, where the sharp trailing-edge morphs into the rounded leading-edge, and vice-versa. A NACA0012 airfoil is approximated in a piecewise linear manner and straight, rigid outer profile links used to define the airfoil contour. The end points of the profile links connect to control links, each set on a central actuation rod via an offset. Chordwise motion of the actuation rod moves the control and the profile links and reverses the airfoil. The paper describes the design methodology and evolution of the final design, based on which two reversible airfoil ribs were fabricated and used to assemble a finite span reversible rotor/wing demonstrator. The profile links were connected by Aluminum strips running in the spanwise direction which provided stiffness as well as support for a pre-tensioned elastomeric skin. An inter-rib connector with a curved-front nose piece supports the leading-edge. The model functioned well and was able to reverse smoothly back-and-forth, on application and reversal of a voltage to the motor. Navier-Stokes CFD simulations (using the TURNS code) show that the drag coefficient of the reversible airfoil (which had a 13% maximum thickness due to the thickness of the profile links) was comparable to that of the NACA0013 airfoil. The drag of a 16% thick elliptical airfoil was, on average, about twice as large, while that of a NACA0012 in reverse flow was 4-5 times as large, even prior to stall. The maximum lift coefficient of the reversible airfoil was lower than the elliptical airfoil, but higher than the NACA0012 in reverse flow operation.
Approximation concepts for numerical airfoil optimization
NASA Technical Reports Server (NTRS)
Vanderplaats, G. N.
1979-01-01
An efficient algorithm for airfoil optimization is presented. The algorithm utilizes approximation concepts to reduce the number of aerodynamic analyses required to reach the optimum design. Examples are presented and compared with previous results. Optimization efficiency improvements of more than a factor of 2 are demonstrated. Improvements in efficiency are demonstrated when analysis data obtained in previous designs are utilized. The method is a general optimization procedure and is not limited to this application. The method is intended for application to a wide range of engineering design problems.
Turbine airfoil having outboard and inboard sections
Mazzola, Stefan; Marra, John J
2015-03-17
A turbine airfoil usable in a turbine engine and formed from at least an outboard section and an inboard section such that an inner end of the outboard section is attached to an outer end of the inboard section. The outboard section may be configured to provide a tip having adequate thickness and may extend radially inward from the tip with a generally constant cross-sectional area. The inboard section may be configured with a tapered cross-sectional area to support the outboard section.
Direct Numerical Simulations of Plunging Airfoils
2010-01-07
Schmidt and E Turkel, Numerical Solutions of the Euler Equations by Finite Volume Methods Using Runge-Kutta Time-Stepping Schemes, AIAA paper 81-1259...Ω ( p ∂vj ∂xj − σij ∂v i ∂xj ) dV (4) Definition 1 A numerical scheme to solve the viscous Navier-Stokes equations is said to be Kinetic Energy...Direct Numerical Simulations of Plunging Airfoils Yves Allaneau∗ and Antony Jameson† Stanford University, Stanford, California, 94305, USA This paper
Feasibility of Actively Cooled Silicon Nitride Airfoil for Turbine Applications Demonstrated
NASA Technical Reports Server (NTRS)
Bhatt, Ramakrishna T.
2001-01-01
Nickel-base superalloys currently limit gas turbine engine performance. Active cooling has extended the temperature range of service of nickel-base superalloys in current gas turbine engines, but the margin for further improvement appears modest. Therefore, significant advancements in materials technology are needed to raise turbine inlet temperatures above 2400 F to increase engine specific thrust and operating efficiency. Because of their low density and high-temperature strength and thermal conductivity, in situ toughened silicon nitride ceramics have received a great deal of attention for cooled structures. However, the high processing costs and low impact resistance of silicon nitride ceramics have proven to be major obstacles for widespread applications. Advanced rapid prototyping technology in combination with conventional gel casting and sintering can reduce high processing costs and may offer an affordable manufacturing approach. Researchers at the NASA Glenn Research Center, in cooperation with a local university and an aerospace company, are developing actively cooled and functionally graded ceramic structures. The objective of this program is to develop cost-effective manufacturing technology and experimental and analytical capabilities for environmentally stable, aerodynamically efficient, foreign-object-damage-resistant, in situ toughened silicon nitride turbine nozzle vanes, and to test these vanes under simulated engine conditions. Starting with computer aided design (CAD) files of an airfoil and a flat plate with internal cooling passages, the permanent and removable mold components for gel casting ceramic slips were made by stereolithography and Sanders machines, respectively. The gel-cast part was dried and sintered to final shape. Several in situ toughened silicon nitride generic airfoils with internal cooling passages have been fabricated. The uncoated and thermal barrier coated airfoils and flat plates were burner rig tested for 30 min without
Parameterized Linear Longitudinal Airship Model
NASA Technical Reports Server (NTRS)
Kulczycki, Eric; Elfes, Alberto; Bayard, David; Quadrelli, Marco; Johnson, Joseph
2010-01-01
A parameterized linear mathematical model of the longitudinal dynamics of an airship is undergoing development. This model is intended to be used in designing control systems for future airships that would operate in the atmospheres of Earth and remote planets. Heretofore, the development of linearized models of the longitudinal dynamics of airships has been costly in that it has been necessary to perform extensive flight testing and to use system-identification techniques to construct models that fit the flight-test data. The present model is a generic one that can be relatively easily specialized to approximate the dynamics of specific airships at specific operating points, without need for further system identification, and with significantly less flight testing. The approach taken in the present development is to merge the linearized dynamical equations of an airship with techniques for estimation of aircraft stability derivatives, and to thereby make it possible to construct a linearized dynamical model of the longitudinal dynamics of a specific airship from geometric and aerodynamic data pertaining to that airship. (It is also planned to develop a model of the lateral dynamics by use of the same methods.) All of the aerodynamic data needed to construct the model of a specific airship can be obtained from wind-tunnel testing and computational fluid dynamics
Control of Shortwave Radiation Parameterization on Tropical Climate Simulation
NASA Astrophysics Data System (ADS)
Crétat, J.; Masson, S. G.; Berthet, S.; Samson, G.; Terray, P.; Dudhia, J.; Pinsard, F.; Hourdin, C.
2015-12-01
SST-forced tropical-channel simulations are used to quantify the control of shortwave (SW) parameterization on the mean tropical climate compared to other major model settings (convection, boundary layer turbulence, vertical and horizontal resolutions). The physical mechanisms whereby this control manifests are explored by the means of a large set of simulations with two widely used SW schemes. Analyses focus on the spatial distribution and magnitude of the net SW radiation budget at the surface (SWnet_SFC), latent heat fluxes, and rainfall at the annual timescale. The model skill and sensitivity to the settings tested are quantified relative to observations and reanalyses and using an ensemble approach. Model skill is mainly controlled by SW parameterization, especially the magnitude of SWnet_SFC and rainfall and both the spatial distribution and magnitude of latent heat fluxes over ocean. On the other hand, the spatial distribution of continental rainfall (SWnet_SFC) is mainly influenced by convection parameterization and horizontal resolution (boundary layer parameterization and orography). Physical understanding of both the control of SW parameterization and sensitivity to SW schemes is addressed by analyzing the thermal structure of the atmosphere and conducting sensitivity experiments to O3 absorption and SW scattering coefficient. SW parameterization shapes the stability of the atmosphere in two different ways according to whether surface is coupled to atmosphere or not, while O3 absorption has minor effects in our simulations. Over SST-prescribed regions, increasing the amount of SW absorption warms the atmosphere only because surface temperatures are fixed, resulting in increased atmospheric stability. Over surface-atmosphere coupled regions (i.e., land points in our simulations), increasing SW absorption warms both atmospheric and surface temperatures, leading to a shift towards a warmer state and a more intense hydrological cycle. This turns in reversal
Control of flow separation in airfoil/wing design applications
NASA Technical Reports Server (NTRS)
Gally, Thomas A.
1994-01-01
Existing aerodynamic design methods have generally concentrated on the optimization of airfoil or wing shapes to produce a minimum drag while satisfying some basic constraints such as lift, pitching moment, or thickness. Since the minimization of drag almost always precludes the existence of separated flow, the evaluation and validation of these design methods for their robustness and accuracy when separated flow is present has not been aggressively pursued. However, two new applications for these design tools may be expected to include separated flow and the issues of aerodynamic design with this feature must be addressed. The first application of the aerodynamic design tools is the design of airfoils or wings to provide an optimal performance over a wide range of flight conditions (multipoint design). While the definition of 'optimal performance' in the multipoint setting is currently being hashed out, it is recognized that given a wide enough range of flight conditions, it will not be possible to ensure a minimum drag constraint at all conditions, and in fact some amount of separated flow (presumably small) may have to be allowed at the more demanding flight conditions. Thus a multipoint design method must be tolerant of the existence of separated flow and may include some controls upon its extent. The second application is in the design of wings with extended high speed buffet boundaries of their flight envelopes. Buffet occurs on a wing when regions of flow separation have grown to the extent that their time varying pressures induce possible destructive effects upon the wing structure or adversely effect either the aircraft controllability or the passenger comfort. A conservative approach to the expansion of the buffet flight boundary is to simply expand the flight envelope of nonseparated flow under the assumption that buffet will also thus be alleviated. However, having the ability to design a wing with separated flow and thus to control the location, extent
A framework for understanding drag parameterizations for coral reefs
NASA Astrophysics Data System (ADS)
Rosman, Johanna H.; Hench, James L.
2011-08-01
In a hydrodynamic sense, a coral reef is a complex array of obstacles that exerts a net drag force on water moving over the reef. This drag is typically parameterized in ocean circulation models using drag coefficients (CD) or roughness length scales (z0); however, published CD for coral reefs span two orders of magnitude, posing a challenge to predictive modeling. Here we examine the reasons for the large range in reported CD and assess the limitations of using CD and z0 to parameterize drag on reefs. Using a formal framework based on the 3-D spatially averaged momentum equations, we show that CD and z0 are functions of canopy geometry and velocity profile shape. Using an idealized two-layer model, we illustrate that CD can vary by more than an order of magnitude for the same geometry and flow depending on the reference velocity selected and that differences in definition account for much of the range in reported CD values. Roughness length scales z0 are typically used in 3-D circulation models to adjust CD for reference height, but this relies on spatially averaged near-bottom velocity profiles being logarithmic. Measurements from a shallow backreef indicate that z0 determined from fits to point measurements of velocity profiles can be very different from z0 required to parameterize spatially averaged drag. More sophisticated parameterizations for drag and shear stresses are required to simulate 3-D velocity fields over shallow reefs; in the meantime, we urge caution when using published CD and z0 values for coral reefs.
FLEET Velocimetry Measurements on a Transonic Airfoil
NASA Technical Reports Server (NTRS)
Burns, Ross A.; Danehy, Paul M.
2017-01-01
Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0 deg, 3.5 deg, and 7deg. Measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty in the context of the applied flowfield. Measurement precisions as low as 1 m/s were observed, while overall uncertainties ranged from 4 to 5 percent. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack.
The modelling of symmetric airfoil vortex generators
NASA Technical Reports Server (NTRS)
Reichert, B. A.; Wendt, B. J.
1996-01-01
An experimental study is conducted to determine the dependence of vortex generator geometry and impinging flow conditions on shed vortex circulation and crossplane peak vorticity for one type of vortex generator. The vortex generator is a symmetric airfoil having a NACA 0012 cross-sectional profile. The geometry and flow parameters varied include angle-of-attack alfa, chordlength c, span h, and Mach number M. The vortex generators are mounted either in isolation or in a symmetric counter-rotating array configuration on the inside surface of a straight pipe. The turbulent boundary layer thickness to pipe radius ratio is delta/R = 0. 17. Circulation and peak vorticity data are derived from crossplane velocity measurements conducted at or about 1 chord downstream of the vortex generator trailing edge. Shed vortex circulation is observed to be proportional to M, alfa, and h/delta. With these parameters held constant, circulation is observed to fall off in monotonic fashion with increasing airfoil aspect ratio AR. Shed vortex peak vorticity is also observed to be proportional to M, alfa, and h/delta. Unlike circulation, however, peak vorticity is observed to increase with increasing aspect ratio, reaching a peak value at AR approx. 2.0 before falling off.
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.
Airfoil for a gas turbine engine
Liang, George
2011-05-24
An airfoil is provided for a turbine of a gas turbine engine. The airfoil comprises: an outer structure comprising a first wall including a leading edge, a trailing edge, a pressure side, and a suction side; an inner structure comprising a second wall spaced from the first wall and at least one intermediate wall; and structure extending between the first and second walls so as to define first and second gaps between the first and second walls. The second wall and the at least one intermediate wall define at least one pressure side supply cavity and at least one suction side supply cavity. The second wall may include at least one first opening near the leading edge of the first wall. The first opening may extend from the at least one pressure side supply cavity to the first gap. The second wall may further comprise at least one second opening near the trailing edge of the outer structure. The second opening may extend from the at least one suction side supply cavity to the second gap. The first wall may comprise at least one first exit opening extending from the first gap through the pressure side of the first wall and at least one second exit opening extending from the second gap through the suction side of the second wall.
Efficiency of an auto-propelled flapping airfoil
NASA Astrophysics Data System (ADS)
Benkherouf, T.; Mekadem, M.; Oualli, H.; Hanchi, S.; Keirsbulck, L.; Labraga, L.
2011-05-01
The present study deals with an investigation of the flow aerodynamic characteristics and the propulsive velocity of a system equipped with a nature inspired propulsion system. In particular, the study is aimed at studying the effect of the flapping frequency on the flow behavior. We consider a NACA0014 airfoil undergoing a vertical sinusoidal flapping motion. In contrast to nearly all previous studies in the literature, the present work does not impose any velocity on the inlet flow. During each iteration the outer flow velocity is computed after having determined the forces exerted on the airfoil. Forward motion may only be produced by flapping motion of the airfoil. This is more consistent with the physical phenomenon. The non-stationary viscous flow around the flapping airfoil is simulated using Ansys-Fluent 12.0.7. The airfoil movement is achieved using the deformable mesh technique and an in-house developed User Define Function (UDF). Our results show the influence of flapping frequency and amplitude on both the airfoil velocity and the propulsive efficiency. The resulting motion is contrasts to the applied forces. In the present study, the frequency ranges from 0.1 to 20 Hz while the airfoil amplitude values considered are: 10%, 17.5%, 25% and 40%.
On the Theory of the Unsteady Motion of an Airfoil
NASA Technical Reports Server (NTRS)
Sedov, L. I.
1947-01-01
The paper presents a systematical analysis of the problem of the determination of the unsteady motion about an airfoil moving in an infinite fluid that contains a system of vortices and the determination of the hydrodynamical forces acting on the airfoil. The hydrodynamical problem is reduced to the determination of the function f (xi) which transforms conformally the external region of the airfoil into the interior of a circle. The proposed methods of determining the irrotational motion of a fluid that is produced by any motion of the airfoil are especially simple and effective if the function f (xi) is rational. As an example the flow is determined for the case of an arbitrary motion of an airfoil of the Joukowsky type. The formulas obtained for the determination of the hydrodynamical forces by means of contour integration are similar to those given by S. Chaplygin. These formulas are used to determine the force acting on the airfoil in the cases where the unsteady motion is potential throughout and the circulation about the airfoil is constant and also when the fluid contains a system of vortices. A full discussion is given of the concept of virtual masses together with practical formulas for computing the virtual mass coefficients.
Inverse design of airfoils using a flexible membrane method
NASA Astrophysics Data System (ADS)
Thinsurat, Kamon
The Modified Garabedian Mc-Fadden (MGM) method is used to inversely design airfoils. The Finite Difference Method (FDM) for Non-Uniform Grids was developed to discretize the MGM equation for numerical solving. The Finite Difference Method (FDM) for Non-Uniform Grids has the advantage of being used flexibly with an unstructured grids airfoil. The commercial software FLUENT is being used as the flow solver. Several conditions are set in FLUENT such as subsonic inviscid flow, subsonic viscous flow, transonic inviscid flow, and transonic viscous flow to test the inverse design code for each condition. A moving grid program is used to create a mesh for new airfoils prior to importing meshes into FLUENT for the analysis of flows. For validation, an iterative process is used so the Cp distribution of the initial airfoil, the NACA0011, achieves the Cp distribution of the target airfoil, the NACA2315, for the subsonic inviscid case at M=0.2. Three other cases were carried out to validate the code. After the code validations, the inverse design method was used to design a shock free airfoil in the transonic condition and to design a separation free airfoil at a high angle of attack in the subsonic condition.
Sensitivity Analysis of Chaotic Flow around Two-Dimensional Airfoil
NASA Astrophysics Data System (ADS)
Blonigan, Patrick; Wang, Qiqi; Nielsen, Eric; Diskin, Boris
2015-11-01
Computational methods for sensitivity analysis are invaluable tools for fluid dynamics research and engineering design. These methods are used in many applications, including aerodynamic shape optimization and adaptive grid refinement. However, traditional sensitivity analysis methods, including the adjoint method, break down when applied to long-time averaged quantities in chaotic fluid flow fields, such as high-fidelity turbulence simulations. This break down is due to the ``Butterfly Effect'' the high sensitivity of chaotic dynamical systems to the initial condition. A new sensitivity analysis method developed by the authors, Least Squares Shadowing (LSS), can compute useful and accurate gradients for quantities of interest in chaotic dynamical systems. LSS computes gradients using the ``shadow trajectory'', a phase space trajectory (or solution) for which perturbations to the flow field do not grow exponentially in time. To efficiently compute many gradients for one objective function, we use an adjoint version of LSS. This talk will briefly outline Least Squares Shadowing and demonstrate it on chaotic flow around a Two-Dimensional airfoil.
Conformal Surface Parameterization for Texture Mapping
1999-03-25
Conformal Surface Parameterization for Texture Mapping Steven Haker Department of Electrical and Computer Engineering University of Minnesota...also like to thank Professor Victoria Interrante for some very helpful conversations on texture mappings. References [1] S. Angenent, S. Haker , A
Performance of two transonic airfoil wind tunnels utilizing limited ventilation
NASA Technical Reports Server (NTRS)
Lee, J. D.; Gregorek, G. M.
1984-01-01
A limited-zone ventilated wall panel was developed for a closed-wall icing tunnel which permitted correct simulation of transonic flow over model rotor airfoil sections with and without ice accretions. Candidate porous panels were tested in the Ohio State University 6- x 12-inch transonic airfoil tunnel and result in essentially interference-free flow, as evidenced by pressure distributions over a NACA 0012 airfoil for Mach numbers up to 0.75. Application to the NRC 12- x 12-inch icing tunnel showed a similar result, which allowed proper transonic flow simulation in that tunnel over its full speed range.
Characteristics of an Airfoil as Affected by Fabric Sag
NASA Technical Reports Server (NTRS)
Ward, Kenneth E
1932-01-01
This report presents the results of tests made at a high value of the Reynolds Number in the N.A.C.A. variable-density wind tunnel to determine the aerodynamic characteristics of an airfoil as affected by fabric sag. Tests were made of two Gottingen 387 airfoils, one having the usual smooth surface and the other having a surface modified to simulate two types of fabric sag. The results of these tests indicate that the usual sagging of the wind covering between ribs has a very small effect on the aerodynamic characteristics of an airfoil.
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.
Transonic airfoil calculations using solution-adaptive grids
NASA Technical Reports Server (NTRS)
Holst, T. L.; Brown, D.
1981-01-01
A new algorithm for generating solution-adaptive grids (SAG) about airfoil configurations embedded in transonic flow is presented. The present SAG approach uses only the airfoil surface solution to recluster grid points on the airfoil surface, i.e., the reclustering problem is one dimension smaller than the flow-field calculation problem. Special controls automatically built into the elliptic grid generation procedure are then used to obtain grids with suitable interior behavior. This concept of redistributing grid points greatly simplifies the idea of solution-adaptive grids. Numerical results indicate significant improvements in accuracy for SAG grids relative to standard grids using the same number of points.
Navier-Stokes analysis of blunt trailing edge airfoils
NASA Technical Reports Server (NTRS)
Stanaway, Sharon; Mccroskey, W. J.; Kroo, Ilan
1992-01-01
The flow around blunt trailing edge airfoils was studied by solving the Reynolds-averaged Navier-Stokes equations. The solution procedure combines a grid around the airfoil with a second grid for the wake so that the time advancement over the domain is fully implicit. This is not only very efficient for the algorithm but also allows implicit solutions of a one equation turbulence model appropriate for both boundary layers and wakes. An algebraic and two one-equation turbulence models are tested for a blunt RAE 2822 airfoil section and detailed comparisons with experimental data are presented in the trailing edge region.
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.
Broadband Noise Predictions for an Airfoil in a Turbulent Stream
NASA Technical Reports Server (NTRS)
Casper, J.; Farassat, F.; Mish, P. F.; Devenport, W. J.
2003-01-01
Loading noise is predicted from unsteady surface pressure measurements on a NACA 0015 airfoil immersed in grid-generated turbulence. The time-dependent pressure is obtained from an array of synchronized transducers on the airfoil surface. Far field noise is predicted by using the time-dependent surface pressure as input to Formulation 1A of Farassat, a solution of the Ffowcs Williams - Hawkings equation. Acoustic predictions are performed with and without the effects of airfoil surface curvature. Scaling rules are developed to compare the present far field predictions with acoustic measurements that are available in the literature.
Laser velocimetry measurements of oscillating airfoil dynamic stall flow field
NASA Technical Reports Server (NTRS)
Chandrasekhara, M. S.; Ahmed, S.
1991-01-01
Ensemble-averaged two-component velocity measurements over an airfoil experiencing oscillatory dynamic stall under compressibility conditions were obtained. The measurements show the formation of a separation bubble over the airfoil that persists till angles of attack close to when the dynamic stall vortex forms and convects. The fluid attains mean velocities as large as 1.6 times the free stream velocity with instantaneous values of 1.8 times the free stream velocity. The airfoil motion induces these large velocities in regions that are far removed from the surface.
Design and experimental results for the S809 airfoil
Somers, D M
1997-01-01
A 21-percent-thick, laminar-flow airfoil, the S809, for horizontal-axis wind-turbine applications, has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of restrained maximum lift, insensitive to roughness, and low profile drag have been achieved. The airfoil also exhibits a docile stall. Comparisons of the theoretical and experimental results show good agreement. Comparisons with other airfoils illustrate the restrained maximum lift coefficient as well as the lower profile-drag coefficients, thus confirming the achievement of the primary objectives.
Design and experimental results for the S805 airfoil
Somers, D.M.
1997-01-01
An airfoil for horizontal-axis wind-turbine applications, the S805, has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of restrained maximum lift, insensitive to roughness, and low profile drag have been achieved. The airfoil also exhibits a docile stall. Comparisons of the theoretical and experimental results show good agreement. Comparisons with other airfoils illustrate the restrained maximum lift coefficient as well as the lower profile-drag coefficients, thus confirming the achievement of the primary objectives.
Brain surface conformal parameterization with algebraic functions.
Wang, Yalin; Gu, Xianfeng; Chan, Tony F; Thompson, Paul M; Yau, Shing-Tung
2006-01-01
In medical imaging, parameterized 3D surface models are of great interest for anatomical modeling and visualization, statistical comparisons of anatomy, and surface-based registration and signal processing. Here we introduce a parameterization method based on algebraic functions. By solving the Yamabe equation with the Ricci flow method, we can conformally map a brain surface to a multi-hole disk. The resulting parameterizations do not have any singularities and are intrinsic and stable. To illustrate the technique, we computed parameterizations of several types of anatomical surfaces in MRI scans of the brain, including the hippocampi and the cerebral cortices with various landmark curves labeled. For the cerebral cortical surfaces, we show the parameterization results are consistent with selected landmark curves and can be matched to each other using constrained harmonic maps. Unlike previous planar conformal parameterization methods, our algorithm does not introduce any singularity points. It also offers a method to explicitly match landmark curves between anatomical surfaces such as the cortex, and to compute conformal invariants for statistical comparisons of anatomy.
Multiple parameterization for hydraulic conductivity identification.
Tsai, Frank T-C; Li, Xiaobao
2008-01-01
Hydraulic conductivity identification remains a challenging inverse problem in ground water modeling because of the inherent nonuniqueness and lack of flexibility in parameterization methods. This study introduces maximum weighted log-likelihood estimation (MWLLE) along with multiple generalized parameterization (GP) methods to identify hydraulic conductivity and to address nonuniqueness and inflexibility problems in parameterization. A scaling factor for information criteria is suggested to obtain reasonable weights of parameterization methods for the MWLLE and model averaging method. The scaling factor is a statistical parameter relating to a desired significance level in Occam's window and the variance of the chi-squares distribution of the fitting error. Through model averaging with multiple GP methods, the conditional estimate of hydraulic conductivity and its total conditional covariances are calculated. A numerical example illustrates the issue arising from Occam's window in estimating model weights and shows the usefulness of the scaling factor to obtain reasonable model weights. Moreover, the numerical example demonstrates the advantage of using multiple GP methods over the zonation and interpolation methods because GP provides better models in the model averaging method. The methodology is applied to the Alamitos Gap area, California, to identify the hydraulic conductivity field. The results show that the use of the scaling factor is necessary in order to incorporate good parameterization methods and to avoid a dominant parameterization method.
A bulk cloud parameterization in a Venus General Circulation Model
NASA Astrophysics Data System (ADS)
Lee, Christopher; Lewis, Stephen R.; Read, Peter L.
2010-04-01
A condensing cloud parameterization is included in a super-rotating Venus General Circulation Model. A parameterization including condensation, evaporation and sedimentation of mono-modal sulfuric acid cloud particles is described. Saturation vapor pressure of sulfuric acid vapor is used to determine cloud formation through instantaneous condensation and destruction through evaporation, while pressure dependent viscosity of a carbon dioxide atmosphere is used to determine sedimentation rates assuming particles fall at their terminal Stokes velocity. Modifications are described to account for the large range of the Reynolds number seen in the Venus atmosphere. Two GCM experiments initialized with 10 ppm-equivalent of sulfuric acid are integrated for 30 Earth years and the results are discussed with reference to "Y" shaped cloud structures observed on Venus. The GCM is able to produce an analog of the "Y" shaped cloud structure through dynamical processes alone, with contributions from the mean westward wind, the equatorial Kelvin wave, and the mid-latitude/polar Mixed Rossby/Gravity waves. The cloud top height in the GCM decreases from equator to pole and latitudinal gradients of cloud top height are comparable to those observed by Pioneer Venus and Venus Express, and those produced in more complex microphysical models of the sulfur cycle on Venus. Differences between the modeled cloud structures and observations are described and dynamical explanations are suggested for the most prominent differences.
Study of the TRAC Airfoil Table Computational System
NASA Technical Reports Server (NTRS)
Hu, Hong
1999-01-01
The report documents the study of the application of the TRAC airfoil table computational package (TRACFOIL) to the prediction of 2D airfoil force and moment data over a wide range of angle of attack and Mach number. The TRACFOIL generates the standard C-81 airfoil table for input into rotorcraft comprehensive codes such as CAM- RAD. The existing TRACFOIL computer package is successfully modified to run on Digital alpha workstations and on Cray-C90 supercomputers. A step-by-step instruction for using the package on both computer platforms is provided. Application of the newer version of TRACFOIL is made for two airfoil sections. The C-81 data obtained using the TRACFOIL method are compared with those of wind-tunnel data and results are presented.
Viscous effects on transonic airfoil stability and response
NASA Technical Reports Server (NTRS)
Berry, H. M.; Batina, J. T.; Yang, T. Y.
1985-01-01
Viscous effects on transonic airfoil stability and response are investigated using an integral boundary layer model coupled to the inviscid XTRAN2L transonic small disturbance code. Unsteady transonic airloads required for stability analyses are computed using a pulse transfer function analysis including viscous effects. The pulse analysis provides unsteady aerodynamic forces for a wide range of reduced frequency in a single flow field computation. Nonlinear time marching aeroelastic solutions are presented which show the effects of viscosity on airfoil response behavior and flutter. Effects of amplitude on time marching responses are demonstrated. A state space aeroelastic model employing Pade approximants to describe the unsteady airloads is used to study the effects of viscosity on transonic airfoil stability. State space dynamic pressure root loci are in good overall agreement with time marching damping and frequency estimates. Parallel sets of results with and without viscous effects reveal the effects of viscosity on transonic unsteady airloads and aeroelastic characteristics of airfoils.
First-stage high pressure turbine bucket airfoil
Brown, Theresa A.; Ahmadi, Majid; Clemens, Eugene; Perry, II, Jacob C.; Holiday, Allyn K.; Delehanty, Richard A.; Jacala, Ariel Caesar
2004-05-25
The first-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinates defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.
Technology for pressure-instrumented thin airfoil models
NASA Technical Reports Server (NTRS)
Wigley, David A.
1988-01-01
A novel method of airfoil model construction was developed. This Laminated Sheet technique uses 0.8 mm thick sheets of A286 containing a network of pre-formed channels which are vacuum brazed together to form the airfoil. A 6.25 percent model of the X29A canard, which has a 5 percent thick section, was built using this technique. The model contained a total of 96 pressure orifices, 56 in three chordwise rows on the upper surface and 37 in three similar rows on the lower surface. It was tested in the NASA Langley 0.3 m Transonic Cryogenic Tunnel. Unique aerodynamic data was obtained over the full range of temperature and pressure. Part of the data was at transonic Mach numbers and flight Reynolds number. A larger two dimensional model of the NACA 64a-105 airfoil section was also fabricated. Scale up presented some problems, but a testable airfoil was fabricated.
Aeroacoustics and aerodynamic performance of a rotor with flatback airfoils.
Paquette, Joshua A.; Barone, Matthew Franklin; Christiansen, Monica; Simley, Eric
2010-06-01
The aerodynamic performance and aeroacoustic noise sources of a rotor employing flatback airfoils have been studied in field test campaign and companion modeling effort. The field test measurements of a sub-scale rotor employing nine meter blades include both performance measurements and acoustic measurements. The acoustic measurements are obtained using a 45 microphone beamforming array, enabling identification of both noise source amplitude and position. Semi-empirical models of flatback airfoil blunt trailing edge noise are developed and calibrated using available aeroacoustic wind tunnel test data. The model results and measurements indicate that flatback airfoil noise is less than drive train noise for the current test turbine. It is also demonstrated that the commonly used Brooks, Pope, and Marcolini model for blunt trailing edge noise may be over-conservative in predicting flatback airfoil noise for wind turbine applications.
High-flaps for natural laminar flow airfoils
NASA Technical Reports Server (NTRS)
Morgan, Harry L.
1986-01-01
A review of the NACA and NASA low-drag airfoil research is presented with particular emphasis given to the development of mechanical high-lift flap systems and their application to general aviation aircraft. These flap systems include split, plain, single-slotted, and double-slotted trailing-edge flaps plus slat and Krueger leading-edge devices. The recently developed continuous variable-camber high-lift mechanism is also described. The state-of-the-art of theoretical methods for the design and analysis of multi-component airfoils in two-dimensional subsonic flow is discussed, and a detailed description of the Langley MCARF (Multi-Component Airfoil Analysis Program) computer code is presented. The results of a recent effort to design a single- and double-slotted flap system for the NASA high speed natural laminar flow (HSNLF) (1)-0213 airfoil using the MCARF code are presented to demonstrate the capabilities and limitations of the code.
Design and experimental results for the S814 airfoil
Somers, D.M.
1997-01-01
A 24-percent-thick airfoil, the S814, for the root region of a horizontal-axis wind-turbine blade has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moment and the airfoil thickness have been satisfied. Comparisons of the theoretical and experimental results show good agreement with the exception of maximum lift which is overpredicted. Comparisons with other airfoils illustrate the higher maximum lift and the lower profile drag of the S814 airfoil, thus confirming the achievement of the objectives.
Transonic airfoil and wing design using Navier-Stokes codes
NASA Technical Reports Server (NTRS)
Yu, N. J.; Campbell, R. L.
1992-01-01
An iterative design method has been implemented into 2D and 3D Navier-Stokes codes for the design of airfoils or wings with given target pressure distributions. The method begins with the analysis of an initial geometry, and obtains the analysis pressure distributions of that geometry. The differences between analysis pressures and target pressures are used to drive geometry changes through the use of a streamline curvature method. This paper describes the procedure that makes the iterative design method work for Navier-Stokes codes. Examples of 2D airfoil design, and 3D wing design are included. It is demonstrated that the method is highly effective for airfoil or wing design at flow conditions where no substantial separation occurs. Problems encountered in the airfoil design with shock induced flow separations are discussed.
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
Turbine airfoil with a compliant outer wall
Campbell, Christian X [Oviedo, FL; Morrison, Jay A [Oviedo, FL
2012-04-03
A turbine airfoil usable in a turbine engine with a cooling system and a compliant dual wall configuration configured to enable thermal expansion between inner and outer layers while eliminating stress formation in the outer layer is disclosed. The compliant dual wall configuration may be formed a dual wall formed from inner and outer layers separated by a support structure. The outer layer may be a compliant layer configured such that the outer layer may thermally expand and thereby reduce the stress within the outer layer. The outer layer may be formed from a nonplanar surface configured to thermally expand. In another embodiment, the outer layer may be planar and include a plurality of slots enabling unrestricted thermal expansion in a direction aligned with the outer layer.
Three-dimensional effects on airfoils
NASA Technical Reports Server (NTRS)
Chevallier, J. P.
1983-01-01
The effects of boundary layer flows along the walls of wind tunnels were studied to validate the transfer of two dimensional calculations to three dimensional transonic flowfield calculations. Results from trials in various wind tunnels were examind to determine the effects of the wall boundary flow on the control surfaces of an airfoil. Models sliding along a groove in the wall of a channel at sub- and transonic speeds were examined, with the finding that with either nonuniformities in the groove, or even if the channel walls are uniform, the lateral boundary layer can cause variations in the central flow region or alter the onset of shock at the transition point. Models for the effects in both turbulence and in the absence of turbulence are formulated, and it is noted that the characteristics of individual wind tunnels must be studied to quantify any existing three dimensional effects.
Experimental airfoil characterization under tailored turbulent conditions
NASA Astrophysics Data System (ADS)
Heißelmann, Hendrik; Peinke, Joachim; Hölling, Michael
2016-09-01
Studies of the impact of turbulent inflow conditions on the airfoil characteristics were performed within the EU FP7 project AVATAR. The aim of this study is to provide data for the validation of simulations and the improvement of engineering tools. Chord-wise pressure distributions and highly-resolved force data of the wind turbine dedicated DU 00-W-212 profile were measured in the wind tunnel in two tailored turbulent inflow conditions generated with an active grid. A sinusoidal and an intermittent pattern with customized inflow angle fluctuations were generated providing two significantly different distributions of reduced frequencies. The obtained pressure distributions and polars from the unsteady patterns are compared to the laminar baseline case.
Heat Transfer of Airfoils and Plates
NASA Technical Reports Server (NTRS)
Seibert, Otto
1943-01-01
The few available test data on the heat dissipation of wholly or partly heated airfoil models are compared with the corresponding data for the flat plate as obtained by an extension of Prandtl's momentum theory, with differentiation between laminar and turbulent boundary layer and transitional region between both, the extent and appearance of which depend upon certain critical factors. The satisfactory agreement obtained justifies far-reaching conclusions in respect to other profile forms and arrangements of heated surface areas. The temperature relationship of the material quantities in its effect on the heat dissipation is discussed as far as is possible at tk.e present state of research, and it is shown that the profile drag of heated wing surfaces can increase or decrease with the temperature increase depending upon the momentarily existent structure of the boundary layer.
Cooled airfoil in a turbine engine
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.
Airfoil, platform, and cooling passage measurements on a rotating transonic high-pressure turbine
NASA Astrophysics Data System (ADS)
Nickol, Jeremy B.
An experiment was performed at The Ohio State University Gas Turbine Laboratory for a film-cooled high-pressure turbine stage operating at design-corrected conditions, with variable rotor and aft purge cooling flow rates. Several distinct experimental programs are combined into one experiment and their results are presented. Pressure and temperature measurements in the internal cooling passages that feed the airfoil film cooling are used as boundary conditions in a model that calculates cooling flow rates and blowing ratio out of each individual film cooling hole. The cooling holes on the suction side choke at even the lowest levels of film cooling, ejecting more than twice the coolant as the holes on the pressure side. However, the blowing ratios are very close due to the freestream massflux on the suction side also being almost twice as great. The highest local blowing ratios actually occur close to the airfoil stagnation point as a result of the low freestream massflux conditions. The choking of suction side cooling holes also results in the majority of any additional coolant added to the blade flowing out through the leading edge and pressure side rows. A second focus of this dissertation is the heat transfer on the rotor airfoil, which features uncooled blades and blades with three different shapes of film cooling hole: cylindrical, diffusing fan shape, and a new advanced shape. Shaped cooling holes have previously shown immense promise on simpler geometries, but experimental results for a rotating turbine have not previously been published in the open literature. Significant improvement from the uncooled case is observed for all shapes of cooling holes, but the improvement from the round to more advanced shapes is seen to be relatively minor. The reduction in relative effectiveness is likely due to the engine-representative secondary flow field interfering with the cooling flow mechanics in the freestream, and may also be caused by shocks and other
Computation of turbulent near wake for asymmetric airfoils
NASA Technical Reports Server (NTRS)
Deiwert, G. S.
1979-01-01
A numerical procedure for studying the turbulent near wake of two dimensional airfoil sections is presented. The Reynolds Navier-Stokes equations were written for flow about bodies of arbitrary geometry and solved on an arbitrary nonuniform curvilinear computational mesh. Eddy viscosity and Reynolds stress turbulence transport models are considered. Specific examples are shown for airfoil section by using an algebraic viscosity model with streamwise relaxation and the interactive Reynolds stress model.
Design and Experimental Results for the S414 Airfoil
2010-08-01
of most current general-aviation aircraft, including busi - ness jets , as well as unmanned aerial vehicles and all sailplanes. It does, however...RDECOM TR 10-D-112 U.S. ARMY RESEARCH, DEVELOPMENT AND ENGINEERING COMMAND TITLE: Design and Experimental Results for the S414 Airfoil AUTHOR: Dan M...Somers and Mark D. Maughmer COMPANY NAME: Airfoils , Incorporated COMPANY ADDRESS: 122 Rose Drive Port Matilda PA 16870-7535 DATE: August 2010 FINAL
Flow Visualization of Dynamic Stall on an Oscillating Airfoil
1989-09-01
Dynamic Stall; Dynamic lift, ’Unsteady lift; Helicopter retreating blade stall; Oscillating airfoil ; Flow visualization,’Schlieren method ;k ez.S-,’ .0...the degree of MASTER OF SCIENCE IN AERONAUTICAL ENGINEERING from the NAVAL POSTGRADUATE SCHOOL September 1989 Author...and moment behavior is quite different from the static stall associated with fixed-wing airfoils . Helicopter retreating blade stall is a dynamic
Active aerodynamic control of wake-airfoil interaction noise - Experiment
NASA Astrophysics Data System (ADS)
Simonich, J. C.; Lavrich, P. L.; Sofrin, T. G.; Topol, D. A.
A proof of concept experiment is conducted that shows the potential for active aerodynamic control of rotor wake/stator interaction noise in a simplified manner. A single airfoil model representing the stator was fitted with a moveable trailing edge flap controlled by a servo motor. The control system moves the motor driven flap in the correct angular displacement phase and rate to reduce the unsteady load on the airfoil during the wake interaction.
Natural laminar flow airfoil analysis and trade studies
NASA Technical Reports Server (NTRS)
1979-01-01
An analysis of an airfoil for a large commercial transport cruising at Mach 0.8 and the use of advanced computer techniques to perform the analysis are described. Incorporation of the airfoil into a natural laminar flow transport configuration is addressed and a comparison of fuel requirements and operating costs between the natural laminar flow transport and an equivalent turbulent flow transport is addressed.
TRANSEP: A program for high lift separated flow about airfoils
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1980-01-01
A method and program called TRANSEP is presented that can be used for the analysis of the flow about a low speed airfoil under high lift, massive separation conditions. Since the present program is a modification of the direct-inverse TRANDES code, it can also be used for the design and analysis of transonic airfoils, including the effects of weak viscous interaction. Interactions on program usage, program modifications to convert TRANDES to TRANSEP, and sample cases and results are given.
NASA Astrophysics Data System (ADS)
Dehghan Manshadi, Mojtaba; Rabani, Ramin
2016-10-01
Shock formation due to flow compressibility and its interaction with boundary layers has adverse effects on aerodynamic characteristics, such as drag increase and flow separation. The objective of this paper is to appraise the practicability of weakening shock waves and, hence, reducing the wave drag in transonic flight regime using a two-dimensional jagged wall and thereby to gain an appropriate jagged wall shape for future empirical study. Different shapes of the jagged wall, including rectangular, circular, and triangular shapes, were employed. The numerical method was validated by experimental and numerical studies involving transonic flow over the NACA0012 airfoil, and the results presented here closely match previous experimental and numerical results. The impact of parameters, including shape and the length-to-spacing ratio of a jagged wall, was studied on aerodynamic forces and flow field. The results revealed that applying a jagged wall method on the upper surface of an airfoil changes the shock structure significantly and disintegrates it, which in turn leads to a decrease in wave drag. It was also found that the maximum drag coefficient decrease of around 17 % occurs with a triangular shape, while the maximum increase in aerodynamic efficiency (lift-to-drag ratio) of around 10 % happens with a rectangular shape at an angle of attack of 2.26°.
SmaggIce 2D Version 1.8: Software Toolkit Developed for Aerodynamic Simulation Over Iced Airfoils
NASA Technical Reports Server (NTRS)
Choo, Yung K.; Vickerman, Mary B.
2005-01-01
SmaggIce 2D version 1.8 is a software toolkit developed at the NASA Glenn Research Center that consists of tools for modeling the geometry of and generating the grids for clean and iced airfoils. Plans call for the completed SmaggIce 2D version 2.0 to streamline the entire aerodynamic simulation process--the characterization and modeling of ice shapes, grid generation, and flow simulation--and to be closely coupled with the public-domain application flow solver, WIND. Grid generated using version 1.8, however, can be used by other flow solvers. SmaggIce 2D will help researchers and engineers study the effects of ice accretion on airfoil performance, which is difficult to do with existing software tools because of complex ice shapes. Using SmaggIce 2D, when fully developed, to simulate flow over an iced airfoil will help to reduce the cost of performing flight and wind-tunnel tests for certifying aircraft in natural and simulated icing conditions.
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.
Analysis of crossover between local and massive separation on airfoils
NASA Technical Reports Server (NTRS)
Barnett, Mark
1987-01-01
The occurrence of massive separation on airfoils operating at high Reynolds number poses an important problem to the aerodynamicist. In the present study, the phenomenon of crossover, induced by airfoil thickness, between local separation and massive separation is investigated for low speed (incompressible), symmetric flow past realistic airfoil geometries. This problem is studied both for the infinite Reynolds number asymptotic limit using triple-deck theory and for finite Reynolds number using interacting boundary-layer theory. Numerical results are presented which illustrate how the flow evolves from local to massive separation as the airfoil thickness is increased. The results of the triple-deck and the interacting boundary-layer analyses are found to be in qualitative agreement for the NACA four digit series and an uncambered supercritical airfoil. The effect of turbulence on the evolution of the flow is also considered. Solutions are presented for turbulent flows past a NACA 0014 airfoil and a circular cylinder. For the latter case, the calculated surface pressure distribution is found to agree well with experimental data if the proper eddy pressure level is specified.
Symmetric airfoil geometry effects on leading edge noise.
Gill, James; Zhang, X; Joseph, P
2013-10-01
Computational aeroacoustic methods are applied to the modeling of noise due to interactions between gusts and the leading edge of real symmetric airfoils. Single frequency harmonic gusts are interacted with various airfoil geometries at zero angle of attack. The effects of airfoil thickness and leading edge radius on noise are investigated systematically and independently for the first time, at higher frequencies than previously used in computational methods. Increases in both leading edge radius and thickness are found to reduce the predicted noise. This noise reduction effect becomes greater with increasing frequency and Mach number. The dominant noise reduction mechanism for airfoils with real geometry is found to be related to the leading edge stagnation region. It is shown that accurate leading edge noise predictions can be made when assuming an inviscid meanflow, but that it is not valid to assume a uniform meanflow. Analytic flat plate predictions are found to over-predict the noise due to a NACA 0002 airfoil by up to 3 dB at high frequencies. The accuracy of analytic flat plate solutions can be expected to decrease with increasing airfoil thickness, leading edge radius, gust frequency, and Mach number.
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.
Incidence angle effects on convected gust airfoil noise
NASA Technical Reports Server (NTRS)
Kerschen, E. J.; Myers, M. R.
1983-01-01
An analysis is developed which predicts the influence of airfoil mean loading on noise generation due to convected gusts. The theory is based on a linearization of the exact inviscid equations about a nonuniform compressible mean flow and the solution is developed using singular perturbation techniques. The case of a flat plate airfoil, at incidence angle alpha, interacting with three-dimensional disturbances is analyzed. It is found that in the vicinity of the airfoil leading and trailing edges, local regions are present which scale on the disturbance wavelength, with the noise generation concentrated in these regions. Away from the airfoil edges, the mean flow variation is found to be slow compared to the disturbance wavelength and no significant noise generation occurs. The mean flow variation near the leading edge generates additional noise by distorting the convected gust. The cumulative effect of the airfoil mean loading in the trailing edge region produces a 0(1) phase shift between the disturbances on the upper and lower surfaces of the airfoil. A corresponding 0(1) decrease, compared to the alpha = 0 case, is found in the noise generated at the trailing edge.
NASA Technical Reports Server (NTRS)
Thompson, J. F.; Turner, L., III; Long, W. S.; Bearden, J. H.
1979-01-01
The development of a numerical simulation of time dependent, turbulent, compressible flow about two dimensional multi-element airfoils of arbitrary shape is described. The basis of this simulation is a technique of automatic numerical generation of coordinate systems fitted to the multiple bodies regardless of their number or shape. Procedures developed whereby the coordinate lines are automatically concentrated in the boundary layer at any Reynolds number are discussed. The compressible turbulent solution involves an algebraic eddy viscosity turbulence model. The laminar version was run for transonic flow at free stream Mach numbers up to 0.9.
NASA Technical Reports Server (NTRS)
Bragg, Michael B.
1993-01-01
This is the second volume of a report documenting the effect of simulated ice accretion on the aerodynamic performance of a NACA 0012 airfoil. Both an experimentally measured and a computer generated ice shape are studied. The purpose of this report is to present the results of the measurements, not an analysis of the data. Surface pressure, integrated lift and pitching moment data are presented as well as drag from a wake survey. A split hot film probe was used to document the flow-field about the airfoil with simulated ice. Data in the separation bubbles, reattached boundary layer and wake are presented. Both tabulated and graphical data are presented in the paper. The data are also available on computer disk for easy access.
Transonic flow over airfoils at energy input and with account for the real properties of the air
NASA Astrophysics Data System (ADS)
Aul'Chenko, S. M.; Zamuraev, V. P.; Kalinina, A. P.
2010-07-01
We have investigated the influence of molecular (thermodynamic and transfer) properties of the air on the qualitative and quantitative characteristics of the gas-dynamic effects of pulsed-periodic energy input in transonic flow over airfoils. The influence of the transfer properties (viscosity) has been taken into account approximately within the boundary layer. It has been shown that qualitative estimates of the above effects do not depend on the account of molecular properties of the air, but the existence of internal degrees of freedom leads to a marked lowering of the temperature level as compared to the ideal gas model. Account of the viscosity somewhat decreases the energy input estimates. The influence of the airfoil shape on the aerodynamic characteristics at energy input has been investigated.
Parameterization of continental boundary layer clouds
NASA Astrophysics Data System (ADS)
Zhu, Ping; Zhao, Wei
2008-05-01
Large eddy simulations (LESs) of continental boundary layer clouds (BLCs) observed at the southern Great Plains (SGP) are used to study issues associated with the parameterization of sub-grid BLCs in large scale models. It is found that liquid water potential temperature θl and total specific humidity qt, which are often used as parameterization predictors in statistical cloud schemes, do not share the same probability distribution in the cloud layer with θl skewed to the left (negatively skewed) and qt skewed to the right (positively skewed). The skewness and kurtosis change substantially in time and space when the development of continental BLCs undergoes a distinct diurnal variation. The wide range of skewness and kurtosis of θl and qt can hardly be described by a single probability distribution function. To extend the application of the statistical cloud parameterization approach, this paper proposes an innovative cloud parameterization scheme that uses the boundary layer height and the lifting condensation level as the primary parameterization predictors. The LES results indicate that the probability distribution of these two quantities is relatively stable compared with that of θl and qt during the diurnal variation and nearly follows a Gaussian function. Verifications using LES output and the observations collected at the Atmospheric Radiation Measurement (ARM) Climate Research Facility (ARCF) SGP site indicate that the proposed scheme works well to represent continental BLCs.
Brain surface parameterization using Riemann surface structure.
Wang, Yalin; Gu, Xianfeng; Hayashi, Kiralee M; Chan, Tony F; Thompson, Paul M; Yau, Shing-Tung
2005-01-01
We develop a general approach that uses holomorphic 1-forms to parameterize anatomical surfaces with complex (possibly branching) topology. Rather than evolve the surface geometry to a plane or sphere, we instead use the fact that all orientable surfaces are Riemann surfaces and admit conformal structures, which induce special curvilinear coordinate systems on the surfaces. Based on Riemann surface structure, we can then canonically partition the surface into patches. Each of these patches can be conformally mapped to a parallelogram. The resulting surface subdivision and the parameterizations of the components are intrinsic and stable. To illustrate the technique, we computed conformal structures for several types of anatomical surfaces in MRI scans of the brain, including the cortex, hippocampus, and lateral ventricles. We found that the resulting parameterizations were consistent across subjects, even for branching structures such as the ventricles, which are otherwise difficult to parameterize. Compared with other variational approaches based on surface inflation, our technique works on surfaces with arbitrary complexity while guaranteeing minimal distortion in the parameterization. It also offers a way to explicitly match landmark curves in anatomical surfaces such as the cortex, providing a surface-based framework to compare anatomy statistically and to generate grids on surfaces for PDE-based signal processing.
NASA Astrophysics Data System (ADS)
Wang, Yingjun; Benson, David J.
2016-12-01
In this paper, an approach based on the fast point-in-polygon (PIP) algorithm and trimmed elements is proposed for isogeometric topology optimization (TO) with arbitrary geometric constraints. The isogeometric parameterized level-set-based TO method, which directly uses the non-uniform rational basis splines (NURBS) for both level set function (LSF) parameterization and objective function calculation, provides higher accuracy and efficiency than previous methods. The integration of trimmed elements is completed by the efficient quadrature rule that can design the quadrature points and weights for arbitrary geometric shape. Numerical examples demonstrate the efficiency and flexibility of the method.
POET: Parameterized Optimization for Empirical Tuning
Yi, Q; Seymour, K; You, H; Vuduc, R; Quinlan, D
2007-01-29
The excessive complexity of both machine architectures and applications have made it difficult for compilers to statically model and predict application behavior. This observation motivates the recent interest in performance tuning using empirical techniques. We present a new embedded scripting language, POET (Parameterized Optimization for Empirical Tuning), for parameterizing complex code transformations so that they can be empirically tuned. The POET language aims to significantly improve the generality, flexibility, and efficiency of existing empirical tuning systems. We have used the language to parameterize and to empirically tune three loop optimizations-interchange, blocking, and unrolling-for two linear algebra kernels. We show experimentally that the time required to tune these optimizations using POET, which does not require any program analysis, is significantly shorter than that when using a full compiler-based source-code optimizer which performs sophisticated program analysis and optimizations.
A parameterization of cloud droplet nucleation
Ghan, S.J.; Chuang, C.C.; Penner, J.E.
1994-01-01
Droplet nucleation is a fundamental cloud process. The number of aerosols activated to form cloud droplets influences not only the number of aerosols scavenged by clouds but also the size of the cloud droplets. Cloud droplet size influences the cloud albedo and the conversion of cloud water to precipitation. Global aerosol models are presently being developed with the intention of coupling with global atmospheric circulation models to evaluate the influence of aerosols and aerosol-cloud interactions on climate. If these and other coupled models are to address issues of aerosol-interactions, the droplet nucleation process must be adequately represented. Ghan et al. have introduced a droplet nucleation parameterization for a single aerosol type that offers certain advantages over the popular Twomey parameterization. Here we describe the generalization of that parameterization to the case of multiple aerosol types, with estimation of aerosol mass as well as number activated.
Control theory based airfoil design for potential flow and a finite volume discretization
NASA Technical Reports Server (NTRS)
Reuther, J.; Jameson, A.
1994-01-01
This paper describes the implementation of optimization techniques based on control theory for airfoil design. In previous studies it was shown that control theory could be used to devise an effective optimization procedure for two-dimensional profiles in which the shape is determined by a conformal transformation from a unit circle, and the control is the mapping function. The goal of our present work is to develop a method which does not depend on conformal mapping, so that it can be extended to treat three-dimensional problems. Therefore, we have developed a method which can address arbitrary geometric shapes through the use of a finite volume method to discretize the potential flow equation. Here the control law serves to provide computationally inexpensive gradient information to a standard numerical optimization method. Results are presented, where both target speed distributions and minimum drag are used as objective functions.
Order-Sorted Parameterization and Induction
NASA Astrophysics Data System (ADS)
Meseguer, José
Parameterization is one of the most powerful features to make specifications and declarative programs modular and reusable, and our best hope for scaling up formal verification efforts. This paper studies order-sorted parameterization at three different levels: (i) its mathematical semantics; (ii) its operational semantics by term rewriting; and (iii) the inductive reasoning principles that can soundly be used to prove properties about such specifications. It shows that achieving the desired properties at each of these three levels is a considerably subtler matter than for many-sorted specifications, but that such properties can be attained under reasonable conditions.
Musial, W.D.; Butterfield, C.P.; Jenks, M.D.
1990-02-01
At the Solar Energy Research Institute (SERI), we carried out tests to measure the effects of leading-edge roughness on an S809 airfoil using a 10-m, three-bladed, horizontal-axis wind turbine (HAWT). The rotor employed a constant-chord (.457 m) blade geometry with zero twist. Blade structural loads were measured with strain gages mounted at 9 spanwise locations. Airfoil pressure measurements were taken at the 80% spanwise station using 32 pressure taps distributed around the airfoil surface. Detailed inflow measurements were taken using nine R.M. Young Model 8002 propvane anemometers on a vertical plane array (VPA) located 10 m upwind of the test turbine in the prevailing wind direction. The major objective of this test was to determine the sensitivity of the S809 airfoil to roughness on a rotating wind turbine blade. We examined this effect by comparing several parameters. We compared power curves to show the sensitivity of whole rotor performance to roughness. We used pressure measurements to generate pressure distributions at the 80% span which operates at a Reynolds number (Re) of 800,000. We then integrated these distributions to determine the effect of roughness on the section's lift and pressure-drag coefficients. We also used the shapes of these distributions to understand how roughness affects the aerodynamic forces on the airfoil. We also compared rough and smooth wind tunnel data to the rotating blade data to study the effects of blade rotation on the aerodynamic behavior of the airfoil below, near, and beyond stall. 13 refs., 11 figs.
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.
On the general theory of thin airfoils for nonuniform motion
NASA Technical Reports Server (NTRS)
Reissner, Eric
1944-01-01
General thin-airfoil theory for a compressible fluid is formulated as boundary problem for the velocity potential, without recourse to the theory of vortex motion. On the basis of this formulation the integral equation of lifting-surface theory for an incompressible fluid is derived with the chordwise component of the fluid velocity at the airfoil as the function to be determined. It is shown how by integration by parts this integral equation can be transformed into the Biot-Savart theorem. A clarification is gained regarding the use of principal value definitions for the integral which occur. The integral equation of lifting-surface theory is used a s the starting point for the establishment of a theory for the nonstationary airfoil which is a generalization of lifting-line theory for the stationary airfoil and which might be called "lifting-strip" theory. Explicit expressions are given for section lift and section moment in terms of the circulation function, which for any given wing deflection is to be determined from an integral equation which is of the type of the equation of lifting-line theory. The results obtained are for airfoils of uniform chord. They can be extended to tapered airfoils. One of the main uses of the results should be that they furnish a practical means for the analysis of the aerodynamic span effect in the problem of wing flutter. The range of applicability of "lifting-strip" theory is the same as that of lifting-line theory so that its results may be applied to airfoils with aspect ratios as low as three.
Assessment of Ice Shape Roughness Using a Self-Orgainizing Map Approach
NASA Technical Reports Server (NTRS)
Mcclain, Stephen T.; Kreeger, Richard E.
2013-01-01
Self-organizing maps are neural-network techniques for representing noisy, multidimensional data aligned along a lower-dimensional and nonlinear manifold. For a large set of noisy data, each element of a finite set of codebook vectors is iteratively moved in the direction of the data closest to the winner codebook vector. Through successive iterations, the codebook vectors begin to align with the trends of the higher-dimensional data. Prior investigations of ice shapes have focused on using self-organizing maps to characterize mean ice forms. The Icing Research Branch has recently acquired a high resolution three dimensional scanner system capable of resolving ice shape surface roughness. A method is presented for the evaluation of surface roughness variations using high-resolution surface scans based on a self-organizing map representation of the mean ice shape. The new method is demonstrated for 1) an 18-in. NACA 23012 airfoil 2 AOA just after the initial ice coverage of the leading 5 of the suction surface of the airfoil, 2) a 21-in. NACA 0012 at 0AOA following coverage of the leading 10 of the airfoil surface, and 3) a cold-soaked 21-in.NACA 0012 airfoil without ice. The SOM method resulted in descriptions of the statistical coverage limits and a quantitative representation of early stages of ice roughness formation on the airfoils. Limitations of the SOM method are explored, and the uncertainty limits of the method are investigated using the non-iced NACA 0012 airfoil measurements.
Approaches for Subgrid Parameterization: Does Scaling Help?
NASA Astrophysics Data System (ADS)
Yano, Jun-Ichi
2016-04-01
Arguably the scaling behavior is a well-established fact in many geophysical systems. There are already many theoretical studies elucidating this issue. However, the scaling law is slow to be introduced in "operational" geophysical modelling, notably for weather forecast as well as climate projection models. The main purpose of this presentation is to ask why, and try to answer this question. As a reference point, the presentation reviews the three major approaches for traditional subgrid parameterization: moment, PDF (probability density function), and mode decomposition. The moment expansion is a standard method for describing the subgrid-scale turbulent flows both in the atmosphere and the oceans. The PDF approach is intuitively appealing as it directly deals with a distribution of variables in subgrid scale in a more direct manner. The third category, originally proposed by Aubry et al (1988) in context of the wall boundary-layer turbulence, is specifically designed to represent coherencies in compact manner by a low--dimensional dynamical system. Their original proposal adopts the proper orthogonal decomposition (POD, or empirical orthogonal functions, EOF) as their mode-decomposition basis. However, the methodology can easily be generalized into any decomposition basis. The mass-flux formulation that is currently adopted in majority of atmospheric models for parameterizing convection can also be considered a special case of the mode decomposition, adopting the segmentally-constant modes for the expansion basis. The mode decomposition can, furthermore, be re-interpreted as a type of Galarkin approach for numerically modelling the subgrid-scale processes. Simple extrapolation of this re-interpretation further suggests us that the subgrid parameterization problem may be re-interpreted as a type of mesh-refinement problem in numerical modelling. We furthermore see a link between the subgrid parameterization and downscaling problems along this line. The mode
Two experimental supercritical laminar-flow-control swept-wing airfoils
NASA Technical Reports Server (NTRS)
Allison, Dennis O.; Dagenhart, J. Ray
1987-01-01
Two supercritical laminar-flow-control airfoils were designed for a large-chord swept-wing experiment in the Langley 8-Foot Transonic Pressure Tunnel where suction was provided through most of the model surface for boundary-layer control. The first airfoil was derived from an existing full-chord laminar airfoil by extending the trailing edge and making changes in the two lower-surface concave regions. The second airfoil differed from the first one in that it was designed for testing without suction in the forward concave region of the lower surface. Differences between the first airfoil and the one from which it was derived as well as between the first and second airfoils are discussed. Airfoil coordinates and predicted pressure distributions for the design normal Mach number of 0.755 and section lift coefficient of 0.55 are given for the three airfoils.
The impact of forest architecture parameterization on GPP simulations
NASA Astrophysics Data System (ADS)
Firanj, Ana; Lalic, Branislava; Podrascanin, Zorica
2015-08-01
The presence of a forest strongly affects ecosystem fluxes by acting as a source or sink of mass and energy. The objective of this study was to investigate the influence of the vertical forest heterogeneity parameterization on gross primary production (GPP) simulations. To introduce a heterogeneity effect, a new method for the upscaling of the leaf level GPP is proposed. This upscaling method is based on the relationship between the leaf area index ( LAI) and the leaf area density ( LAD) profiles and the standard sun/shade leaf separation method. The effect of the crown shape and foliage distribution parameterization on the simulated GPP is confirmed in a comparison study between the proposed method and the standard sun/shade upscaling method. The observed values used in the comparison study are assimilated during the vegetation period on three distinguished forest eddy-covariance (EC) measurement sites chosen for the diversity of their morphological characteristics. The obtained results show (a) the sensitivity of the simulated GPP to the leaf area density profile, (b) the capability of the proposed scaling method to calculate the contribution of the different canopy layers to the entire canopy GPP, and (c) a better agreement with the observations of the simulated GPP with the proposed upscaling method compared with the standard sun/shade method.
Finite Difference Calculation of an Inviscid Transonic Flow over Oscillating Airfoils,
1980-10-01
solutions for the NLR 7301 airfoil and the NACA 64A010 airfoil also will be obtained. They will be compared with the results obtained by Tijdeman and...and the NACA 0012 airfoil oscillating in pitch, in order to obtain several individual flow patterns. The resulting unsteady pressure distributions...shock wave locations, etc, are presented. Furthermore, the unsteady numerical results obtained by this procedure for the NLR 7301 airfoil and the NACA
Modified-Dewan Optical Turbulence Parameterizations
2007-11-02
Kea Observatories on the Island of Hawaii (Businger et al. 2002) by converting standard Numerical Weather Prediction (NWP) forecast model output into...describing optical turbulence. The Dewan parameterization is also being used to forecast optical seeing conditions for ground-based telescopes at the Mauna
Parameterization guidelines and considerations for hydrologic models
Technology Transfer Automated Retrieval System (TEKTRAN)
Imparting knowledge of the physical processes of a system to a model and determining a set of parameter values for a hydrologic or water quality model application (i.e., parameterization) is an important and difficult task. An exponential increase in literature has been devoted to the use and develo...
Parameterizing cloud condensation nuclei concentrations during HOPE
NASA Astrophysics Data System (ADS)
Hande, Luke B.; Engler, Christa; Hoose, Corinna; Tegen, Ina
2016-09-01
An aerosol model was used to simulate the generation and transport of aerosols over Germany during the HD(CP)2 Observational Prototype Experiment (HOPE) field campaign of 2013. The aerosol number concentrations and size distributions were evaluated against observations, which shows satisfactory agreement in the magnitude and temporal variability of the main aerosol contributors to cloud condensation nuclei (CCN) concentrations. From the modelled aerosol number concentrations, number concentrations of CCN were calculated as a function of vertical velocity using a comprehensive aerosol activation scheme which takes into account the influence of aerosol chemical and physical properties on CCN formation. There is a large amount of spatial variability in aerosol concentrations; however the resulting CCN concentrations vary significantly less over the domain. Temporal variability is large in both aerosols and CCN. A parameterization of the CCN number concentrations is developed for use in models. The technique involves defining a number of best fit functions to capture the dependence of CCN on vertical velocity at different pressure levels. In this way, aerosol chemical and physical properties as well as thermodynamic conditions are taken into account in the new CCN parameterization. A comparison between the parameterization and the CCN estimates from the model data shows excellent agreement. This parameterization may be used in other regions and time periods with a similar aerosol load; furthermore, the technique demonstrated here may be employed in regions dominated by different aerosol species.
Uncertainty Analysis for a Jet Flap Airfoil
NASA Technical Reports Server (NTRS)
Green, Lawrence L.; Cruz, Josue
2006-01-01
An analysis of variance (ANOVA) study was performed to quantify the potential uncertainties of lift and pitching moment coefficient calculations from a computational fluid dynamics code, relative to an experiment, for a jet flap airfoil configuration. Uncertainties due to a number of factors including grid density, angle of attack and jet flap blowing coefficient were examined. The ANOVA software produced a numerical model of the input coefficient data, as functions of the selected factors, to a user-specified order (linear, 2-factor interference, quadratic, or cubic). Residuals between the model and actual data were also produced at each of the input conditions, and uncertainty confidence intervals (in the form of Least Significant Differences or LSD) for experimental, computational, and combined experimental / computational data sets were computed. The LSD bars indicate the smallest resolvable differences in the functional values (lift or pitching moment coefficient) attributable solely to changes in independent variable, given just the input data points from selected data sets. The software also provided a collection of diagnostics which evaluate the suitability of the input data set for use within the ANOVA process, and which examine the behavior of the resultant data, possibly suggesting transformations which should be applied to the data to reduce the LSD. The results illustrate some of the key features of, and results from, the uncertainty analysis studies, including the use of both numerical (continuous) and categorical (discrete) factors, the effects of the number and range of the input data points, and the effects of the number of factors considered simultaneously.
Numerical solution of periodic vortical flows about a thin airfoil
NASA Technical Reports Server (NTRS)
Scott, James R.; Atassi, Hafiz M.
1989-01-01
A numerical method is developed for computing periodic, three-dimensional, vortical flows around isolated airfoils. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Solutions for thin airfoils at zero degrees incidence to the mean flow are presented in this paper. Using an elliptic coordinate transformation, the computational domain is transformed into a rectangle. The Sommerfeld radiation condition is applied to the unsteady pressure on the grid line corresponding to the far field boundary. The results are compared with a Possio solver, and it is shown that for maximum accuracy the grid should depend on both the Mach number and reduced frequency. Finally, in order to assess the range of validity of the classical thin airfoil approximation, results for airfoils with zero thickness are compared with results for airfoils with small thickness.
Computational Analysis of Dual Radius Circulation Control Airfoils
NASA Technical Reports Server (NTRS)
Lee-Rausch, E. M.; Vatsa, V. N.; Rumsey, C. L.
2006-01-01
The goal of the work is to use multiple codes and multiple configurations to provide an assessment of the capability of RANS solvers to predict circulation control dual radius airfoil performance and also to identify key issues associated with the computational predictions of these configurations that can result in discrepancies in the predicted solutions. Solutions were obtained for the Georgia Tech Research Institute (GTRI) dual radius circulation control airfoil and the General Aviation Circulation Control (GACC) dual radius airfoil. For the GTRI-DR airfoil, two-dimensional structured and unstructured grid computations predicted the experimental trend in sectional lift variation with blowing coefficient very well. Good code to code comparisons between the chordwise surface pressure coefficients and the solution streamtraces also indicated that the detailed flow characteristics were matched between the computations. For the GACC-DR airfoil, two-dimensional structured and unstructured grid computations predicted the sectional lift and chordwise pressure distributions accurately at the no blowing condition. However at a moderate blowing coefficient, although the code to code variation was small, the differences between the computations and experiment were significant. Computations were made to investigate the sensitivity of the sectional lift and pressure distributions to some of the experimental and computational parameters, but none of these could entirely account for the differences in the experimental and computational results. Thus, CFD may indeed be adequate as a prediction tool for dual radius CC flows, but limited and difficult to obtain two-dimensional experimental data prevents a confident assessment at this time.
Supercritical flow past a symmetrical bicircular arc airfoil
NASA Technical Reports Server (NTRS)
Holt, Maurice; Yew, Khoy Chuah
1989-01-01
A numerical scheme is developed for computing steady supercritical flow about symmetrical airfoils, applying it to an ellipse for zero angle of attack. An algorithmic description of this new scheme is presented. Application to a symmetrical bicircular arc airfoil is also proposed. The flow field before the shock is region 1. For transonic flow, singularity can be avoided by integrating the resulting ordinary differential equations away from the body. Region 2 contains the shock which will be located by shock fitting techniques. The shock divides region 2 into supersonic and subsonic regions and there is no singularity problem in this case. The Method of Lines is used in this region and it is advantageous to integrate the resulting ordinary differential equation along the body for shock fitting. Coaxial coordinates have to be used for the bicircular arc airfoil so that boundary values on the airfoil body can be taken with one direction of the coaxial coordinates fixed. To avoid taking boundary values at + or - infinity in the coaxial co-ordinary system, approximate analytical representation of the flow field near the tips of the airfoil is proposed.
Design analysis of vertical wind turbine with airfoil variation
NASA Astrophysics Data System (ADS)
Maulana, Muhammad Ilham; Qaedy, T. Masykur Al; Nawawi, Muhammad
2016-03-01
With an ever increasing electrical energy crisis occurring in the Banda Aceh City, it will be important to investigate alternative methods of generating power in ways different than fossil fuels. In fact, one of the biggest sources of energy in Aceh is wind energy. It can be harnessed not only by big corporations but also by individuals using Vertical Axis Wind Turbines (VAWT). This paper presents a three-dimensional CFD analysis of the influence of airfoil design on performance of a Darrieus-type vertical-axis wind turbine (VAWT). The main objective of this paper is to develop an airfoil design for NACA 63-series vertical axis wind turbine, for average wind velocity 2,5 m/s. To utilize both lift and drag force, some of designs of airfoil are analyzed using a commercial computational fluid dynamics solver such us Fluent. Simulation is performed for this airfoil at different angles of attach rearranging from -12°, -8°, -4°, 0°, 4°, 8°, and 12°. The analysis showed that the significant enhancement in value of lift coefficient for airfoil NACA 63-series is occurred for NACA 63-412.
Empirical parameterization of setup, swash, and runup
Stockdon, H.F.; Holman, R.A.; Howd, P.A.; Sallenger, A.H.
2006-01-01
Using shoreline water-level time series collected during 10 dynamically diverse field experiments, an empirical parameterization for extreme runup, defined by the 2% exceedence value, has been developed for use on natural beaches over a wide range of conditions. Runup, the height of discrete water-level maxima, depends on two dynamically different processes; time-averaged wave setup and total swash excursion, each of which is parameterized separately. Setup at the shoreline was best parameterized using a dimensional form of the more common Iribarren-based setup expression that includes foreshore beach slope, offshore wave height, and deep-water wavelength. Significant swash can be decomposed into the incident and infragravity frequency bands. Incident swash is also best parameterized using a dimensional form of the Iribarren-based expression. Infragravity swash is best modeled dimensionally using offshore wave height and wavelength and shows no statistically significant linear dependence on either foreshore or surf-zone slope. On infragravity-dominated dissipative beaches, the magnitudes of both setup and swash, modeling both incident and infragravity frequency components together, are dependent only on offshore wave height and wavelength. Statistics of predicted runup averaged over all sites indicate a - 17 cm bias and an rms error of 38 cm: the mean observed runup elevation for all experiments was 144 cm. On intermediate and reflective beaches with complex foreshore topography, the use of an alongshore-averaged beach slope in practical applications of the runup parameterization may result in a relative runup error equal to 51% of the fractional variability between the measured and the averaged slope.
2013-12-24
helicopter rotor blades, wind turbine blades, pitching and flapping airfoils and wings , and rotating turbomachinery blades. For instance, helicopter...of turbulent flow over a pitching airfoil at realistic Reynolds and Mach numbers is performed. Numerical stability at high Reynolds number...Approved for Public Release; Distribution Unlimited Large-Eddy Simulation Analysis of Unsteady Separation Over a Pitching Airfoil at High Reynolds
Evaluation of Airfoil Dynamic Stall Characteristics for Maneuverability
NASA Technical Reports Server (NTRS)
Bousman, William G.; Aiken, Edwin W. (Technical Monitor)
2000-01-01
In severe maneuvers, out of necessity for a military aircraft or inadvertently for a civil aircraft, a helicopter airfoil will stall in a dynamic manner and provide lift beyond what would be calculated based on static airfoil tests. The augmented lift that occurs in dynamic stall is related to a vortex that is shed near the leading edge of the airfoil. However, directly related to the augmented lift that results from the dynamic stall vortex are significant penalties in pitching moment and drag. An understanding of the relationship between the augmented lift in dynamic stall and the associated moment and drag penalties is the purpose of this paper. This relationship is characterized using data obtained in two-dimensional wind tunnel tests and related to the problem of helicopter maneuverability.
Ice Accretions on a Swept GLC-305 Airfoil
NASA Technical Reports Server (NTRS)
Vargas, Mario; Papadakis, Michael; Potapczuk, Mark; Addy, Harold; Sheldon, David; Giriunas, Julius
2002-01-01
An experiment was conducted in the Icing Research Tunnel (IRT) at NASA Glenn Research Center to obtain castings of ice accretions formed on a 28 deg. swept GLC-305 airfoil that is representative of a modern business aircraft wing. Because of the complexity of the casting process, the airfoil was designed with three removable leading edges covering the whole span. Ice accretions were obtained at six icing conditions. After the ice was accreted, the leading edges were detached from the airfoil and moved to a cold room. Molds of the ice accretions were obtained, and from them, urethane castings were fabricated. This experiment is the icing test of a two-part experiment to study the aerodynamic effects of ice accretions.
Inverse boundary-layer technique for airfoil design
NASA Technical Reports Server (NTRS)
Henderson, M. L.
1979-01-01
A description is presented of a technique for the optimization of airfoil pressure distributions using an interactive inverse boundary-layer program. This program allows the user to determine quickly a near-optimum subsonic pressure distribution which meets his requirements for lift, drag, and pitching moment at the desired flow conditions. The method employs an inverse turbulent boundary-layer scheme for definition of the turbulent recovery portion of the pressure distribution. Two levels of pressure-distribution architecture are used - a simple roof top for preliminary studies and a more complex four-region architecture for a more refined design. A technique is employed to avoid the specification of pressure distributions which result in unrealistic airfoils, that is, those with negative thickness. The program allows rapid evaluation of a designed pressure distribution off-design in Reynolds number, transition location, and angle of attack, and will compute an airfoil contour for the designed pressure distribution using linear theory.
Effects of finite aspect ratio on wind turbine airfoil measurements
NASA Astrophysics Data System (ADS)
Kiefer, Janik; Miller, Mark A.; Hultmark, Marcus; Hansen, Martin O. L.
2016-09-01
Wind turbines partly operate in stalled conditions within their operational cycle. To simulate these conditions, it is also necessary to obtain 2-D airfoil data in terms of lift and drag coefficients at high angles of attack. Such data has been obtained previously, but often at low aspect ratios and only barely past the stall point, where strong wall boundary layer influence is expected. In this study, the influence of the wall boundary layer on 2D airfoil data, especially in the post stall domain, is investigated. Here, a wind turbine airfoil is tested at different angles of attack and with two aspect ratios of AR = 1 and AR = 2. The tests are conducted in a wind tunnel that is pressurized up to 150 bar in order to achieve a constant Reynolds number of Rec = 3 • 106, despite the variable chord length.
Aerodynamic performance of an annular classical airfoil cascade
NASA Technical Reports Server (NTRS)
Bergsten, D. E.; Stauter, R. C.; Fleeter, S.
1983-01-01
Results are presented for a series of experiments that were performed in a large-scale subsonic annular cascade facility that was specifically designed to provide three-dimensional aerodynamic data for the verification of numerical-calculation codes. In particular, the detailed three-dimensional aerodynamic performance of a classical flat-plate airfoil cascade is determined for angles of incidence of 0, 5, and 10 deg. The resulting data are analyzed and are correlated with predictions obtained from NASA's MERIDL and TSONIC numerical programs. It is found that: (1) at 0 and 5 deg, the airfoil surface data show a good correlation with the predictions; (2) at 10 deg, the data are in fair agreement with the numerical predictions; and (3) the two-dimensional Gaussian similarity relationship is appropriate for the wake velocity profiles in the mid-span region of the airfoil.
Computation of viscous transonic flow about a lifting airfoil
NASA Technical Reports Server (NTRS)
Walitt, L.; Liu, C. Y.
1976-01-01
The viscous transonic flow about a stationary body in free air was numerically investigated. The geometry chosen was a symmetric NACA 64A010 airfoil at a freestream Mach number of 0.8, a Reynolds number of 4 million based on chord, and angles of attack of 0 and 2 degrees. These conditions were such that, at 2 degrees incidence unsteady periodic motion was calculated along the aft portion of the airfoil and in its wake. Although no unsteady measurements were made for the NACA 64A010 airfoil at these flow conditions, interpolated steady measurements of lift, drag, and surface static pressures compared favorably with corresponding computed time-averaged lift, drag, and surface static pressures.
A Simple Parameterization of 3 x 3 Magic Squares
ERIC Educational Resources Information Center
Trenkler, Gotz; Schmidt, Karsten; Trenkler, Dietrich
2012-01-01
In this article a new parameterization of magic squares of order three is presented. This parameterization permits an easy computation of their inverses, eigenvalues, eigenvectors and adjoints. Some attention is paid to the Luoshu, one of the oldest magic squares.
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.
European upper mantle tomography: adaptively parameterized models
NASA Astrophysics Data System (ADS)
Schäfer, J.; Boschi, L.
2009-04-01
We have devised a new algorithm for upper-mantle surface-wave tomography based on adaptive parameterization: i.e. the size of each parameterization pixel depends on the local density of seismic data coverage. The advantage in using this kind of parameterization is that a high resolution can be achieved in regions with dense data coverage while a lower (and cheaper) resolution is kept in regions with low coverage. This way, parameterization is everywhere optimal, both in terms of its computational cost, and of model resolution. This is especially important for data sets with inhomogenous data coverage, as it is usually the case for global seismic databases. The data set we use has an especially good coverage around Switzerland and over central Europe. We focus on periods from 35s to 150s. The final goal of the project is to determine a new model of seismic velocities for the upper mantle underlying Europe and the Mediterranean Basin, of resolution higher than what is currently found in the literature. Our inversions involve regularization via norm and roughness minimization, and this in turn requires that discrete norm and roughness operators associated with our adaptive grid be precisely defined. The discretization of the roughness damping operator in the case of adaptive parameterizations is not as trivial as it is for the uniform ones; important complications arise from the significant lateral variations in the size of pixels. We chose to first define the roughness operator in a spherical harmonic framework, and subsequently translate it to discrete pixels via a linear transformation. Since the smallest pixels we allow in our parameterization have a size of 0.625 °, the spherical-harmonic roughness operator has to be defined up to harmonic degree 899, corresponding to 810.000 harmonic coefficients. This results in considerable computational costs: we conduct the harmonic-pixel transformations on a small Beowulf cluster. We validate our implementation of adaptive
NASA Technical Reports Server (NTRS)
1979-01-01
A comprehensive review of all NASA airfoil research, conducted both in-house and under grant and contract, as well as a broad spectrum of airfoil research outside of NASA is presented. Emphasis is placed on the development of computational aerodynamic codes for airfoil analysis and design, the development of experimental facilities and test techniques, and all types of airfoil applications.
Airfoil design by numerical optimization using a minicomputer
NASA Technical Reports Server (NTRS)
Hicks, R. M.; Szelazek, C. A.
1978-01-01
A computer program developed for the automated design of low speed airfoils utilizes a generalized Joukowski method for aerodynamic analysis coupled with a conjugate gradient, penalty function, numerical optimization algorithm to give an efficient calculation technique for use with minicomputers. The program designs airfoils with a prescribed pressure distribution as well as those which minimize or maximize some aerodynamic force coefficient. At present the method is restricted to inviscid, incompressible flow. A typical design problem will execute in 4.5 hr on an HP 9830 minicomputer.
Euler computations of AGARD Working Group 07 airfoil test cases
NASA Technical Reports Server (NTRS)
Pulliam, T. H.; Barton, J. T.
1985-01-01
In an attempt to provide a set of accurate standard test problems for computational code developers, a series of inviscid airfoil test cases were chosen by the AGARD Working Group 07, a subpanel of the AGARD Fluid Dynamics Panel. The cases include three different airfoils at transonic to supersonic conditions. A large number of international experts responded in this effort with computations that have been contrasted for accuracy and consistency. This paper is a summary of the authors' contribution to this study. In particular, the important aspects of the solution process that made it possible to obtain the high level of accuracy needed in this study are stressed.
Differential pressure sensing system for airfoils usable in turbine engines
Yang, Wen-Ching; Stampahar, Maria E.
2005-09-13
A detection system for identifying airfoils having a cooling systems with orifices that are plugged with contaminants or with showerheads having a portion burned off. The detection system measures pressures at different locations and calculates or measures a differential pressure. The differential pressure may be compared with a known benchmark value to determine whether the differential pressure has changed. Changes in the differential pressure may indicate that one or more of the orifices in a cooling system of an airfoil are plugged or that portions of, or all of, a showerhead has burned off.
Evolving aerodynamic airfoils for wind turbines through a genetic algorithm
NASA Astrophysics Data System (ADS)
Hernández, J. J.; Gómez, E.; Grageda, J. I.; Couder, C.; Solís, A.; Hanotel, C. L.; Ledesma, JI
2017-01-01
Nowadays, genetic algorithms stand out for airfoil optimisation, due to the virtues of mutation and crossing-over techniques. In this work we propose a genetic algorithm with arithmetic crossover rules. The optimisation criteria are taken to be the maximisation of both aerodynamic efficiency and lift coefficient, while minimising drag coefficient. Such algorithm shows greatly improvements in computational costs, as well as a high performance by obtaining optimised airfoils for Mexico City's specific wind conditions from generic wind turbines designed for higher Reynolds numbers, in few iterations.
Dynamic Stall of a Pitching and Horizontally Oscillating Airfoil
NASA Astrophysics Data System (ADS)
Martinat, G.; Braza#, M.; Harran, G.; Sevrain, A.; Tzabiras, G.; Hoarau, Y.; Favier, D.
This paper provides a study of the dynamic stall of a pitching airfoil and of a pitching and horizontally oscillating airfoil at 105 Reynolds number by means of nbumerical simulation. Three turbulence models are compared in both cases: URANS Spalart-Allmaras model, URANS k—ɛ Chien model and URANS=OES model. results are in accordance with experimental data but spalart model seems to be too much viscous to provide good results and overpredict hysteresis cycle observed where URANS=OES seems to be viscousless. URANS k—ɛ Chien model is providing the best results.
Feasibility of predicting performance degradation of airfoils in heavy rain
NASA Technical Reports Server (NTRS)
Bilanin, A. J.; Quackenbush, T. R.; Feo, A.
1989-01-01
The heavy rain aerodynamic performance penalty program is detailed. This effort supported the design of a fullscale test program as well as examined the feasibility of estimating the degradation of performance of airfoils from first principles. The analytic efforts were supplemented by a droplet splashback test program in an attempt to observe the physics of impact and generation of ejecta. These tests demonstrated that the interaction of rain with an airfoil is a highly complex phenomenon and this interaction is not likely to be analyzed analytically with existing tools.
A hybrid algorithm for transonic airfoil and wing design
NASA Technical Reports Server (NTRS)
Campbell, Richard L.; Smith, Leigh A.
1987-01-01
The present method for the design of transonic airfoils and wings employs a predictor/corrector approach in which an analysis code calculates the flowfield for an initial geometry, then modifies it on the basis of the difference between calculated and target pressures. This allows the design method to be straightforwardly coupled with any existing analysis code, as presently undertaken with several two- and three-dimensional potential flow codes. The results obtained indicate that the method is robust and accurate, even in the cases of airfoils with strongly supercritical flow and shocks. The design codes are noted to require computational resources typical of current pure-inverse methods.
Modeling of heavy-gas effects on airfoil flows
NASA Technical Reports Server (NTRS)
Drela, Mark
1992-01-01
Thermodynamic models were constructed for a calorically imperfect gas and for a non-ideal gas. These were incorporated into a quasi one dimensional flow solver to develop an understanding of the differences in flow behavior between the new models and the perfect gas model. The models were also incorporated into a two dimensional flow solver to investigate their effects on transonic airfoil flows. Specifically, the calculations simulated airfoil testing in a proposed high Reynolds number heavy gas test facility. The results indicate that the non-idealities caused significant differences in the flow field, but that matching of an appropriate non-dimensional parameter led to flows similar to those in air.
Aspects of Numerical Simulation of Circulation Control Airfoils
NASA Technical Reports Server (NTRS)
Swanson, R. C.; Rumsey, C. L.; Anders, S. G.
2005-01-01
The mass-averaged compressible Navier-Stokes equations are solved for circulation control airfoils. Numerical solutions are computed with a multigrid method that uses an implicit approximate factorization smoother. The effects of flow conditions (e.g., free-stream Mach number, angle of attack, momentum coefficient) and mesh on the prediction of circulation control airfoil flows are considered. In addition, the impact of turbulence modeling, including curvature effects and modifications to reduce eddy viscosity levels in the wall jet (i.e., Coanda flow), is discussed. Computed pressure distributions are compared with available experimental data.
Turbulent Mixing Parameterizations for Oceanic Flows and Student Support
2014-09-30
projects is to formulate robust turbulence parameterizations that are applicable for a wide range of oceanic flow conditions. OBJECTIVES The...primary objectives of these projects are to bridge the gap between parameterizations/models for small-scale turbulent mixing developed from fundamental...1 DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited. Turbulent Mixing Parameterizations for Oceanic Flows
Parameterization of cloud effects on the absorption of solar radiation
NASA Technical Reports Server (NTRS)
Davies, R.
1983-01-01
A radiation parameterization for the NASA Goddard climate model was developed, tested, and implemented. Interactive and off-hire experiments with the climate model to determine the limitations of the present parameterization scheme are summarized. The parameterization of Cloud absorption in terms of solar zeith angle, column water vapors about the cloud top, and cloud liquid water content is discussed.
Parameterization of Cumulus Convective Cloud Systems in Mesoscale Forecast Models
2013-09-30
1 DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited. Parameterization of Cumulus Convective Cloud Systems in...parameterization of cumulus convective clouds in mesoscale numerical weather prediction models OBJECTIVES Conduct detailed studies of cloud ...microphysical processes in order to develop a unified parameterization of boundary layer stratocumulus and trade wind cumulus convective clouds . Develop
Effects of Airfoil Thickness and Maximum Lift Coefficient on Roughness Sensitivity: 1997--1998
Somers, D. M.
2005-01-01
A matrix of airfoils has been developed to determine the effects of airfoil thickness and the maximum lift to leading-edge roughness. The matrix consists of three natural-laminar-flow airfoils, the S901, S902, and S903, for wind turbine applications. The airfoils have been designed and analyzed theoretically and verified experimentally in the Pennsylvania State University low-speed, low-turbulence wind tunnel. The effect of roughness on the maximum life increases with increasing airfoil thickness and decreases slightly with increasing maximum lift. Comparisons of the theoretical and experimental results generally show good agreement.
NASA Technical Reports Server (NTRS)
Graham, Donald J
1949-01-01
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 wind-tunnel 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 NACA 6-series airfoils.
NASA Technical Reports Server (NTRS)
Jacobs, Eastman N
1932-01-01
Report presents the results of wind tunnel tests on a group of eight very thick airfoils having sections of the same thickness as those used near the roots of tapered airfoils. The tests were made to study certain discontinuities in the characteristic curves that have been obtained from previous tests of these airfoils, and to compare the characteristics of the different sections at values of the Reynolds number comparable with those attained in flight. The discontinuities were found to disappear as the Reynolds number was increased. The results obtained from the large-scale airfoil, a symmetrical airfoil having a thickness ratio of 21 per cent, has the best general characteristics.
Automatic Parameterization Strategy for Cardiac Electrophysiology Simulations.
Costa, Caroline Mendonca; Hoetzl, Elena; Rocha, Bernardo Martins; Prassl, Anton J; Plank, Gernot
2013-10-01
Driven by recent advances in medical imaging, image segmentation and numerical techniques, computer models of ventricular electrophysiology account for increasingly finer levels of anatomical and biophysical detail. However, considering the large number of model parameters involved parameterization poses a major challenge. A minimum requirement in combined experimental and modeling studies is to achieve good agreement in activation and repolarization sequences between model and experiment or patient data. In this study, we propose basic techniques which aid in determining bidomain parameters to match activation sequences. An iterative parameterization algorithm is implemented which determines appropriate bulk conductivities which yield prescribed velocities. In addition, a method is proposed for splitting the computed bulk conductivities into individual bidomain conductivities by prescribing anisotropy ratios.
Invariant box-parameterization of neutrino oscillations
Weiler, Thomas J.; Wagner, DJ
1998-10-19
The model-independent 'box' parameterization of neutrino oscillations is examined. The invariant boxes are the classical amplitudes of the individual oscillating terms. Being observables, the boxes are independent of the choice of parameterization of the mixing matrix. Emphasis is placed on the relations among the box parameters due to mixing-matrix unitarity, and on the reduction of the number of boxes to the minimum basis set. Using the box algebra, we show that CP-violation may be inferred from measurements of neutrino flavor mixing even when the oscillatory factors have averaged. General analyses of neutrino oscillations among n{>=}3 flavors can readily determine the boxes, which can then be manipulated to yield magnitudes of mixing matrix elements.
Numerical Archetypal Parameterization for Mesoscale Convective Systems
NASA Astrophysics Data System (ADS)
Yano, J. I.
2015-12-01
Vertical shear tends to organize atmospheric moist convection into multiscale coherent structures. Especially, the counter-gradient vertical transport of horizontal momentum by organized convection can enhance the wind shear and transport kinetic energy upscale. However, this process is not represented by traditional parameterizations. The present paper sets the archetypal dynamical models, originally formulated by the second author, into a parameterization context by utilizing a nonhydrostatic anelastic model with segmentally-constant approximation (NAM-SCA). Using a two-dimensional framework as a starting point, NAM-SCA spontaneously generates propagating tropical squall-lines in a sheared environment. A high numerical efficiency is achieved through a novel compression methodology. The numerically-generated archetypes produce vertical profiles of convective momentum transport that are consistent with the analytic archetype.
Parameterizing surface wind speed over complex topography
NASA Astrophysics Data System (ADS)
Helbig, N.; Mott, R.; Herwijnen, A.; Winstral, A.; Jonas, T.
2017-01-01
Subgrid parameterizations are used in coarse-scale meteorological and land surface models to account for the impact of unresolved topography on wind speed. While various parameterizations have been suggested, these were generally validated on a limited number of measurements in specific geographical areas. We used high-resolution wind fields to investigate which terrain parameters most affect near-surface wind speed over complex topography under neutral conditions. Wind fields were simulated using the Advanced Regional Prediction System (ARPS) on Gaussian random fields as model topographies to cover a wide range of terrain characteristics. We computed coarse-scale wind speed, i.e., a spatial average over the large grid cell accounting for influence of unresolved topography, using a previously suggested subgrid parameterization for the sky view factor. We only require correlation length of subgrid topographic features and mean square slope in the coarse grid cell. Computed coarse-scale wind speed compared well with domain-averaged ARPS wind speed. To further statistically downscale coarse-scale wind speed, we use local, fine-scale topographic parameters, namely, the Laplacian of terrain elevations and mean square slope. Both parameters showed large correlations with fine-scale ARPS wind speed. Comparing downscaled numerical weather prediction wind speed with measurements from a large number of stations throughout Switzerland resulted in overall improved correlations and distribution statistics. Since we used a large number of model topographies to derive the subgrid parameterization and the downscaling framework, both are not scale dependent nor bound to a specific geographic region. Both can readily be implemented since they are based on easy to derive terrain parameters.
Aerosol water parameterization: a single parameter framework
NASA Astrophysics Data System (ADS)
Metzger, S.; Steil, B.; Abdelkader, M.; Klingmüller, K.; Xu, L.; Penner, J. E.; Fountoukis, C.; Nenes, A.; Lelieveld, J.
2015-11-01
We introduce a framework to efficiently parameterize the aerosol water uptake for mixtures of semi-volatile and non-volatile compounds, based on the coefficient, νi. This solute specific coefficient was introduced in Metzger et al. (2012) to accurately parameterize the single solution hygroscopic growth, considering the Kelvin effect - accounting for the water uptake of concentrated nanometer sized particles up to dilute solutions, i.e., from the compounds relative humidity of deliquescence (RHD) up to supersaturation (Köhler-theory). Here we extend the νi-parameterization from single to mixed solutions. We evaluate our framework at various levels of complexity, by considering the full gas-liquid-solid partitioning for a comprehensive comparison with reference calculations using the E-AIM, EQUISOLV II, ISORROPIA II models as well as textbook examples. We apply our parameterization in EQSAM4clim, the EQuilibrium Simplified Aerosol Model V4 for climate simulations, implemented in a box model and in the global chemistry-climate model EMAC. Our results show: (i) that the νi-approach enables to analytically solve the entire gas-liquid-solid partitioning and the mixed solution water uptake with sufficient accuracy, (ii) that, e.g., pure ammonium nitrate and mixed ammonium nitrate - ammonium sulfate mixtures can be solved with a simple method, and (iii) that the aerosol optical depth (AOD) simulations are in close agreement with remote sensing observations for the year 2005. Long-term evaluation of the EMAC results based on EQSAM4clim and ISORROPIA II will be presented separately.
Unified Parameterization of the Marine Boundary Layer
2010-09-30
information if it does not display a currently valid OMB control number. 1. REPORT DATE 2010 2 . REPORT TYPE 3. DATES COVERED 00-00-2010 to 00-00-2010...boundary layer closure for the convective boundary layer 2 . An EDMF approach to the vertical transport of TKE in convective boundary layers 3. EDMF in...4 implementation and extension to shallow cumulus parameterization is in progress. 2 An integrated TKE-based eddy-diffusivity/mass-flux
Effect of pivot location and passive heave on propulsion from a pitching airfoil
NASA Astrophysics Data System (ADS)
Mackowski, A. W.; Williamson, C. H. K.
2017-01-01
We experimentally investigate the propulsive characteristics of a pitching NACA 0012 airfoil section, with emphasis on thrust and propulsive efficiency, at a Reynolds number of 1.7 ×104 . For the sake of mechanical simplicity, we consider an airfoil restricted to a single actuator in the pitching direction. We examine the effect of changing the airfoil's axis of rotation, finding that contrary to Garrick's linear theory, there exists a pitching axis near the airfoil that maximizes propulsive efficiency. Next, we examine the effect of placing passive springs on the airfoil in the heave (transverse) direction using our Cyber-Physical Fluid Dynamics technique. This elastic heaving motion allows the airfoil to combine pitching and heaving modes while being actuated only in the pitching direction. Two sets of dynamics are considered: one case where the airfoil is weighted unevenly and pitched about its center of mass (so that the resulting heaving motion is independent of inertial forces), and another case where the airfoil's center of mass is fixed at its centroid. For pitching at an amplitude of 8∘ and a reduced frequency k of two, we find that elastic heave produces a maximum propulsive efficiency of 35%, compared to 25% without any heave motion. Further, while operating at the same efficiency as the static-pivot case, we find that passive heaving greatly increases the magnitude of the airfoil's thrust. The airfoil configurations with highest propulsive efficiency generally involve pitching near or ahead of the airfoil's leading edge.
Aerodynamic shape optimization using preconditioned conjugate gradient methods
NASA Technical Reports Server (NTRS)
Burgreen, Greg W.; Baysal, Oktay
1993-01-01
In an effort to further improve upon the latest advancements made in aerodynamic shape optimization procedures, a systematic study is performed to examine several current solution methodologies as applied to various aspects of the optimization procedure. It is demonstrated that preconditioned conjugate gradient-like methodologies dramatically decrease the computational efforts required for such procedures. The design problem investigated is the shape optimization of the upper and lower surfaces of an initially symmetric (NACA-012) airfoil in inviscid transonic flow and at zero degree angle-of-attack. The complete surface shape is represented using a Bezier-Bernstein polynomial. The present optimization method then automatically obtains supercritical airfoil shapes over a variety of freestream Mach numbers. Furthermore, the best optimization strategy examined resulted in a factor of 8 decrease in computational time as well as a factor of 4 decrease in memory over the most efficient strategies in current use.
Fire parameterization on a global scale
NASA Astrophysics Data System (ADS)
Pechony, O.; Shindell, D. T.
2009-08-01
We present a convenient physically based global-scale fire parameterization algorithm for global climate models. We indicate environmental conditions favorable for fire occurrence based on calculation of the vapor pressure deficit as a function of location and time. Two ignition models are used. One assumes ubiquitous ignition, the other incorporates natural and anthropogenic sources, as well as anthropogenic fire suppression. Evaluation of the method using Global Precipitation Climatology Project precipitation, National Centers for Environmental Prediction/National Center for Atmospheric Research temperature and relative humidity, and Moderate Resolution Imaging Spectroradiometer (MODIS) Leaf Area Index as a proxy for global vegetation density gives results in remarkable correspondence with global fire patterns observed from the MODIS and Visible and Infrared Scanner satellite instruments. The parameterized fires successfully reproduce the spatial distribution of global fires as well as the seasonal variability. The interannual variability of global fire activity derived from the 20-year advanced very high resolution radiometer record are well reproduced using Goddard Institute for Space Studies general circulation models climate simulations, as is the response to the climate changes following the eruptions of El Chichon and Mount Pinatubo. In conjunction with climate models and data sets on vegetation changes with time, the suggested fire parameterization offers the possibility to estimate relative variations of global fire activity for past and future climates.
Thermonuclear Reaction Rate Parameterization for Nuclear Astrophysics
NASA Astrophysics Data System (ADS)
Sharp, Jacob; Kozub, Raymond L.; Smith, Michael S.; Scott, Jason; Lingerfelt, Eric
2004-10-01
The knowledge of thermonuclear reaction rates is vital to simulate novae, supernovae, X-ray bursts, and other astrophysical events. To facilitate dissemination of this knowledge, a set of tools has been created for managing reaction rates, located at www.nucastrodata.org. One tool is a rate parameterizer, which provides a parameterization for nuclear reaction rate vs. temperature values in the most widely used functional form. Currently, the parameterizer uses the Levenberg-Marquardt method (LMM), which requires an initial estimate of the best-fit parameters. The initial estimate is currently provided randomly from a preselected pool. To improve the quality of fits, a new, active method of selecting parameters has been developed. The parameters of each set in the pool are altered for a few iterations to replicate the input data as closely as possible. Then, the set which most nearly matches the input data (based on chi squared) is used in the LMM as the initial estimate for the final fitting procedure. A description of the new, active algorithm and its performance will be presented. Supported by the U. S. Department of Energy.
An Approach to the Constrained Design of Natural Laminar Flow Airfoils
NASA Technical Reports Server (NTRS)
Green, Bradford E.
1997-01-01
A 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. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integral turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the laminar flow toward the desired amount. An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.
An approach to the constrained design of natural laminar flow airfoils
NASA Technical Reports Server (NTRS)
Green, Bradford Earl
1995-01-01
A 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. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integml turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the larninar flow toward the desired amounl An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.
Wind tunnel tests of two airfoils for wind turbines operating at high reynolds numbers
Sommers, D.; Tangler, J.
2000-06-29
The objectives of this study were to verify the predictions of the Eppler Airfoil Design and Analysis Code for Reynolds numbers up to 6 x 106 and to acquire the section characteristics of two airfoils being considered for large, megawatt-size wind turbines. One airfoil, the S825, was designed to achieve a high maximum lift coefficient suitable for variable-speed machines. The other airfoil, the S827, was designed to achieve a low maximum lift coefficient suitable for stall-regulated machines. Both airfoils were tested in the NASA Langley Low-Turbulence Pressure Tunnel (LTPT) for smooth, fixed-transition, and rough surface conditions at Reynolds numbers of 1, 2, 3, 4, and 6 x 106. The results show the maximum lift coefficient of both airfoils is substantially underpredicted for Reynolds numbers over 3 x 106 and emphasized the difficulty of designing low-lift airfoils for high Reynolds numbers.
Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines: Preprint
Migliore, P.; Oerlemans, S.
2003-12-01
Aeroacoustic tests of seven airfoils were performed in an open jet anechoic wind tunnel. Six of the airfoils are candidates for use on small wind turbines operating at low Reynolds number. One airfoil was tested for comparison to benchmark data. Tests were conducted with and without boundary layer tripping. In some cases a turbulence grid was placed upstream in the test section to investigate inflow turbulence noise. An array of 48 microphones was used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing edge noise was dominant for all airfoils in clean tunnel flow. With the boundary layer untripped, several airfoils exhibited pure tones that disappeared after proper tripping was applied. In the presence of inflow turbulence, leading edge noise was dominant for all airfoils.
Control of shortwave radiation parameterization on tropical climate SST-forced simulation
NASA Astrophysics Data System (ADS)
Crétat, Julien; Masson, Sébastien; Berthet, Sarah; Samson, Guillaume; Terray, Pascal; Dudhia, Jimy; Pinsard, Françoise; Hourdin, Christophe
2016-09-01
SST-forced tropical-channel simulations are used to quantify the control of shortwave (SW) parameterization on the mean tropical climate compared to other major model settings (convection, boundary layer turbulence, vertical and horizontal resolutions), and to pinpoint the physical mechanisms whereby this control manifests. Analyses focus on the spatial distribution and magnitude of the net SW radiation budget at the surface (SWnet_SFC), latent heat fluxes, and rainfall at the annual timescale. The model skill and sensitivity to the tested settings are quantified relative to observations and using an ensemble approach. Persistent biases include overestimated SWnet_SFC and too intense hydrological cycle. However, model skill is mainly controlled by SW parameterization, especially the magnitude of SWnet_SFC and rainfall and both the spatial distribution and magnitude of latent heat fluxes over ocean. On the other hand, the spatial distribution of continental rainfall (SWnet_SFC) is mainly influenced by convection parameterization and horizontal resolution (boundary layer parameterization and orography). Physical understanding of the control of SW parameterization is addressed by analyzing the thermal structure of the atmosphere and conducting sensitivity experiments to O3 absorption and SW scattering coefficient. SW parameterization shapes the stability of the atmosphere in two different ways according to whether surface is coupled to atmosphere or not, while O3 absorption has minor effects in our simulations. Over SST-prescribed regions, increasing the amount of SW absorption warms the atmosphere only because surface temperatures are fixed, resulting in increased atmospheric stability. Over land-atmosphere coupled regions, increasing SW absorption warms both atmospheric and surface temperatures, leading to a shift towards a warmer state and a more intense hydrological cycle. This turns in reversal model behavior between land and sea points, with the SW scheme that
NASA Technical Reports Server (NTRS)
Gregorek, Gerald; Dresse, John J.; LaNoe, Karine; Ratvasky, Thomas (Technical Monitor)
2000-01-01
The need for fundamental research in Ice Contaminated Tailplane Stall (ICTS) was established through three international conferences sponsored by the FAA. A joint NASA/FAA Tailplane Icing Program was formed in 1994 with the Ohio State University playing a critical role for wind tunnel and analytical research. Two entries of a full-scale 2-dimensional tailplane airfoil model of a DHC-6 Twin Otter were made in The Ohio State University 7x10 ft wind tunnel. This report describes the second test entry that examined additional ice shapes and roughness, as well as airfoil section differences. The addition data obtained in this test fortified the original database of aerodynamic coefficients that permit a detailed analysis of flight test results with an OSU-developed analytical program. The testing encompassed a full range of angles of attack and elevator deflections at flight Reynolds number conditions. Aerodynamic coefficients, C(L), C(M), and C(He), were obtained by integrating static pressure coefficient, C(P), values obtained from surface taps. Comparisons of clean and iced airfoil results show a significant decrease in the tailplane aeroperformance (decreased C(Lmax), decreased stall angle, increased C(He)) for all ice shapes with the grit having the lease affect and the LEWICE shape having the greatest affect. All results were consistent with observed tailplane stall phenomena and constitute an effective set of data for comprehensive analysis of ICTS.
Effectiveness of Thermal-Pneumatic Airfoil-Ice-Protection System
NASA Technical Reports Server (NTRS)
Gowan, William H., Jr.; Mulholland, Donald R.
1951-01-01
Icing and drag investigations were conducted in the NACA Lewis icing research tunnel employing a combination thermal-pneumatic de-icer mounted on a 42-inch-chord NACA 0018 airfoil. The de-icer consisted of a 3-inch-wide electrically heated strip symmetrically located about the leading edge with inflatable tubes on the upper and lower airfoil surfaces aft of the heated area. The entire de-icer extended to approximately 25 percent of chord. A maximum power density of 9.25 watts per square inch was required for marginal ice protection on the airfoil leading edge at an air temperature of 00 F and an airspeed of 300 miles per hour. Drag measurements indicated, that without icing, the de-icer installation increased the section drag to approximately 140 percent of that of the bare airfoil; with the tubes inflated, this value increased to a maximum of approximately 620 percent. A 2-minute tube-inflation cycle prevented excessive ice formation on the inflatable area although small scattered residual Ice formations remained after inflation and were removed intermittently during later cycles. Effects of the time lag of heater temperatures after initial application of power and the insulating effect of ice formations on heater temperatures were also determined.
Computer Program to Obtain Ordinates for NACA Airfoils
NASA Technical Reports Server (NTRS)
Ladson, Charles L.; Brooks, Cuyler W., Jr.; Hill, Acquilla S.; Sproles, Darrell W.
1996-01-01
Computer programs to produce the ordinates for airfoils of any thickness, thickness distribution, or camber in the NACA airfoil series were developed in the early 1970's and are published as NASA TM X-3069 and TM X-3284. For analytic airfoils, the ordinates are exact. For the 6-series and all but the leading edge of the 6A-series airfoils, agreement between the ordinates obtained from the program and previously published ordinates is generally within 5 x 10(exp -5) chord. Since the publication of these programs, the use of personal computers and individual workstations has proliferated. This report describes a computer program that combines the capabilities of the previously published versions. This program is written in ANSI FORTRAN 77 and can be compiled to run on DOS, UNIX, and VMS based personal computers and workstations as well as mainframes. An effort was made to make all inputs to the program as simple as possible to use and to lead the user through the process by means of a menu.
Cost Effective Repair Techniques for Turbine Airfoils. Volume I
1978-11-01
Turbine blades and vanes in current engines are subjected to the most hostile environment...payoff potential in turbine vanes / blades . The criteria used included: • Incidence of damage - Scrapped or damaged turbine airfoils at the ALC centers...Corporate Author: GENERAL ELECTRIC CO CINCINNATI OHIO AIRCRAFT ENGINE GROUP Unclassified Title: (U) Cost Effective Repair Techniques for Turbine
Cost Effective Repair Techniques for Turbine Airfoils. Volume 2
1979-04-01
BLADES , *GUIDE VANES , *REPAIR, TURBOFAN ENGINES , DIFFUSION BONDING, COST EFFECTIVENESS Identifiers: (U) * Turbine vanes , TF-39 engines , Activated...REPAIR TECHNIQUES FOR TURBINE AIRFOILS J. A. WEIN W. R. YOUNG GENERAL ELECTRIC COMPANY AIRCRAFT ENGINE GROUP CINCINNATI, OHIO 45215 APRIL 1979...Author: GENERAL ELECTRIC CO CINCINNATI OH AIRCRAFT ENGINE BUSINESS GROUP Unclassified Title: (U) Cost Effective Repair Techniques for
Experimental Investigation on Airfoil Shock Control by Plasma Aerodynamic Actuation
NASA Astrophysics Data System (ADS)
Sun, Quan; Cheng, Bangqin; Li, Yinghong; Cui, Wei; Jin, Di; Li, Jun
2013-11-01
An experimental investigation on airfoil (NACA64—215) shock control is performed by plasma aerodynamic actuation in a supersonic tunnel (Ma = 2). The results of schlieren and pressure measurement show that when plasma aerodynamic actuation is applied, the position moves forward and the intensity of shock at the head of the airfoil weakens. With the increase in actuating voltage, the total pressure measured at the head of the airfoil increases, which means that the shock intensity decreases and the control effect increases. The best actuation effect is caused by upwind-direction actuation with a magnetic field, and then downwind-direction actuation with a magnetic field, while the control effect of aerodynamic actuation without a magnetic field is the most inconspicuous. The mean intensity of the normal shock at the head of the airfoil is relatively decreased by 16.33%, and the normal shock intensity is relatively reduced by 27.5% when 1000 V actuating voltage and upwind-direction actuation are applied with a magnetic field. This paper theoretically analyzes the Joule heating effect generated by DC discharge and the Lorentz force effect caused by the magnetic field. The discharge characteristics are compared for all kinds of actuation conditions to reveal the mechanism of shock control by plasma aerodynamic actuation.
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.
Advanced turbine study. [airfoil coling in rocket turbines
NASA Technical Reports Server (NTRS)
1982-01-01
Experiments to determine the available increase in turbine horsepower achieved by increasing turbine inlet temperature over a range of 1800 to 2600 R, while applying current gas turbine airfoil cling technology are discussed. Four cases of rocket turbine operating conditions were investigated. Two of the cases used O2/H2 propellant, one with a fuel flowrate of 160 pps, the other 80 pps. Two cases used O2/CH4 propellant, each having different fuel flowrates, pressure ratios, and inlet pressures. Film cooling was found to be the required scheme for these rocket turbine applications because of the high heat flux environments. Conventional convective or impingement cooling, used in jet engines, is inadequate in a rocket turbine environment because of the resulting high temperature gradients in the airfoil wall, causing high strains and low cyclic life. The hydrogen-rich turbine environment experienced a loss, or no gain, in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The effects of film cooling with regard to reduced flow available for turbine work, dilution of mainstream gas temperature and cooling reentry losses, offset the relatively low specific work capability of hydrogen when increasing turbine inlet temperature over the 1800 to 2600 R range. However, the methane-rich environment experienced an increase in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The results of a materials survey and heat transfer and durability analysis are discussed.
Stability of Inviscid Flow over Airfoils Admitting Multiple Numerical Solutions
NASA Astrophysics Data System (ADS)
Liu, Ya; Xiong, Juntao; Liu, Feng; Luo, Shijun
2012-11-01
Multiple numerical solutions at the same flight condition are found of inviscid transonic flow over certain airfoils (Jameson et al., AIAA 2011-3509) within some Mach number range. Both symmetric and asymmetric solutions exist for a symmetric airfoil at zero angle of attack. Global linear stability analysis of the multiple solutions is conducted. Linear perturbation equations of the Euler equations around a steady-state solution are formed and discretized numerically. An eigenvalue problem is then constructed using the modal analysis approach. Only a small portion of the eigen spectrum is needed and thus can be found efficiently by using Arnoldi's algorithm. The least stable or unstable mode corresponds to the eigenvalue with the largest real part. Analysis of the NACA 0012 airfoil indicates stability of symmetric solutions of the Euler equations at conditions where buffet is found from unsteady Navier-Stokes equations. Euler solutions of the same airfoil but modified to include the displacement thickness of the boundary layer computed from the Navier-Stokes equations, however, exhibit instability based on the present linear stability analysis. Graduate Student.
NASA Astrophysics Data System (ADS)
Zhang, G. Q.; Ji, L. C.
2017-04-01
The vortex formation and shedding process behind an isolated cylinder under the wake interference of an oscillating airfoil at different oscillating frequencies and amplitudes have been studied numerically. The ;2S; and ;P+S; mode are found in the drag-producing von Kármán vortex wake and thrust-producing inverted von Kármán vortex wake, respectively. For the ;2P; case, the pairs of vortices with different sizes are laterally parallel to each other. For the ;2P+2S; case, the merged single vortices do not appear between the vortex pairs, but they only emerge after every two pairs of vortices. For the two-degree of freedom (2-DOF) VIV in the wake of the steady flow, seven types of the phase portraits of the displacement shapes are observed. In the range of small reduced velocity (Vr < 5), the body trajectories are mainly characterized by an ;8-shape; with some differences in phase. At the peak amplitude, the body trajectories have changed into two shapes: i.e., the half ;8-shape; (initial stage) and the crescent-shape (final stage). When Vr = 6 , the body trajectories have also shown to have two shapes: i.e., the ;8-shape; (initial stage) and the half ;8-shape; (final stage). When Vr arrives at the low branch (Vr > 7), the body trajectories will become irregularly shaped like a cone-net.
Aerodynamic Shape Optimization Using Hybridized Differential Evolution
NASA Technical Reports Server (NTRS)
Madavan, Nateri K.
2003-01-01
An aerodynamic shape optimization method that uses an evolutionary algorithm known at Differential Evolution (DE) in conjunction with various hybridization strategies is described. DE is a simple and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems. Various hybridization strategies for DE are explored, including the use of neural networks as well as traditional local search methods. A Navier-Stokes solver is used to evaluate the various intermediate designs and provide inputs to the hybrid DE optimizer. The method is implemented on distributed parallel computers so that new designs can be obtained within reasonable turnaround times. Results are presented for the inverse design of a turbine airfoil from a modern jet engine. (The final paper will include at least one other aerodynamic design application). The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated.
Parameterization of the scattering and absorption properties of individual ice crystals
Yang, Ping; Liou, K. N.; Wyser, Klaus; Mitchell, David
2000-02-27
We present parameterizations of the single-scattering properties for individual ice crystals of various habits based on the results computed from the accurate light scattering calculations. The projected area, volume, and single-scattering properties of ice crystals with various shapes and sizes are computed for 56 narrow spectral bands covering 0.2-5 {mu}m. The ice crystal habits considered in this study are hexagonal plates, solid and hollow columns, planar and spatial bullet rosette, and aggregates that are commonly observed in cirrus clouds. Using the observational relationships between the aspect ratios and the sizes of ice crystals, we can define the three-dimensional structure of these ice crystal habits with respect to their maximum dimensions for light scattering calculations. The volume and projected area of ice crystals, expressed in terms of the diameters of the corresponding equivalent spheres, are first parameterized by employing the ice crystal maximum dimensions. Further, various analytical expressions as functions of the effective dimensions of ice crystals have been developed to parameterize the extinction and absorption efficiencies, asymmetry factor, and the truncation of the forward peak energy in the phase function. The present parameterization scheme provides an efficient approach to obtain the basic scattering and absorption properties of nonspherical ice crystals. (c) 2000 American Geophysical Union.
An experimental study of a bio-inspired corrugated airfoil for micro air vehicle applications
NASA Astrophysics Data System (ADS)
Murphy, Jeffery T.; Hu, Hui
2010-08-01
An experimental study was conducted to investigate the aerodynamic characteristics of a bio-inspired corrugated airfoil compared with a smooth-surfaced airfoil and a flat plate at the chord Reynolds number of Re C = 58,000-125,000 to explore the potential applications of such bio-inspired corrugated airfoils for micro air vehicle designs. In addition to measuring the aerodynamic lift and drag forces acting on the tested airfoils, a digital particle image velocimetry system was used to conduct detailed flowfield measurements to quantify the transient behavior of vortex and turbulent flow structures around the airfoils. The measurement result revealed clearly that the corrugated airfoil has better performance over the smooth-surfaced airfoil and the flat plate in providing higher lift and preventing large-scale flow separation and airfoil stall at low Reynolds numbers (Re C < 100,000). While aerodynamic performance of the smooth-surfaced airfoil and the flat plate would vary considerably with the changing of the chord Reynolds numbers, the aerodynamic performance of the corrugated airfoil was found to be almost insensitive to the Reynolds numbers. The detailed flow field measurements were correlated with the aerodynamic force measurement data to elucidate underlying physics to improve our understanding about how and why the corrugation feature found in dragonfly wings holds aerodynamic advantages for low Reynolds number flight applications.
A Physically Based Fractional Cloudiness Parameterization
1990-07-27
34 just beiow the PBL top. and designed for use as a PBL parameterization in a large-scale an infinitesimal "ventilation layer" just above the Earth s...34convective mass flux" concept the observations of Caughey et al. (1982) and Nicholls and introduced by Arakawa (1969) and adopted in many Turton (1986), who...Nicholls and Turton (1986)]. We can interpret y, as the value of x associated with the downdraft air at level B. Since there is a sharp gradient of V
New Parameterization of Neutron Absorption Cross Sections
NASA Astrophysics Data System (ADS)
Tripathi, Ram K.; Wilson, John W.; Cucinotta, Francis A.
1997-06-01
Recent parameterization of absorption cross sections for any system of charged ion collisions, including proton-nucleus collisions, is extended for neutron-nucleus collisions valid from approx. 1 MeV to a few GeV, thus providing a comprehensive picture of absorption cross sections for any system of collision pairs (charged or uncharged). The parameters are associated with the physics of the problem. At lower energies, optical potential at the surface is important, and the Pauli operator plays an increasingly important role at intermediate energies. The agreement between the calculated and experimental data is better than earlier published results.
New Parameterization of Neutron Absorption Cross Sections
NASA Technical Reports Server (NTRS)
Tripathi, Ram K.; Wilson, John W.; Cucinotta, Francis A.
1997-01-01
Recent parameterization of absorption cross sections for any system of charged ion collisions, including proton-nucleus collisions, is extended for neutron-nucleus collisions valid from approx. 1 MeV to a few GeV, thus providing a comprehensive picture of absorption cross sections for any system of collision pairs (charged or uncharged). The parameters are associated with the physics of the problem. At lower energies, optical potential at the surface is important, and the Pauli operator plays an increasingly important role at intermediate energies. The agreement between the calculated and experimental data is better than earlier published results.
Lightning parameterization in a storm electrification model
NASA Technical Reports Server (NTRS)
Helsdon, John H., Jr.; Farley, Richard D.; Wu, Gang
1988-01-01
The parameterization of an intracloud lightning discharge has been implemented in our Storm Electrification Model. The initiation, propagation direction, termination and charge redistribution of the discharge are approximated assuming overall charge neutrality. Various simulations involving differing amounts of charge transferred have been done. The effects of the lightning-produced ions on the hydrometeor charges, electric field components and electrical energy depend strongly on the charge transferred. A comparison between the measured electric field change of an actual intracloud flash and the field change due to the simulated discharge show favorable agreement.
Hydraulic Conductivity Estimation using Bayesian Model Averaging and Generalized Parameterization
NASA Astrophysics Data System (ADS)
Tsai, F. T.; Li, X.
2006-12-01
Non-uniqueness in parameterization scheme is an inherent problem in groundwater inverse modeling due to limited data. To cope with the non-uniqueness problem of parameterization, we introduce a Bayesian Model Averaging (BMA) method to integrate a set of selected parameterization methods. The estimation uncertainty in BMA includes the uncertainty in individual parameterization methods as the within-parameterization variance and the uncertainty from using different parameterization methods as the between-parameterization variance. Moreover, the generalized parameterization (GP) method is considered in the geostatistical framework in this study. The GP method aims at increasing the flexibility of parameterization through the combination of a zonation structure and an interpolation method. The use of BMP with GP avoids over-confidence in a single parameterization method. A normalized least-squares estimation (NLSE) is adopted to calculate the posterior probability for each GP. We employee the adjoint state method for the sensitivity analysis on the weighting coefficients in the GP method. The adjoint state method is also applied to the NLSE problem. The proposed methodology is implemented to the Alamitos Barrier Project (ABP) in California, where the spatially distributed hydraulic conductivity is estimated. The optimal weighting coefficients embedded in GP are identified through the maximum likelihood estimation (MLE) where the misfits between the observed and calculated groundwater heads are minimized. The conditional mean and conditional variance of the estimated hydraulic conductivity distribution using BMA are obtained to assess the estimation uncertainty.
Wind-tunnel Tests of the NACA 45-125 Airfoil: A Thick Airfoil for High-Speed Airplanes
NASA Technical Reports Server (NTRS)
Delano, James B.
1940-01-01
Investigations of the pressure distribution, the profile drag, and the location of transition for a 30-inch-chord 25-percent-thick N.A,C.A. 45-125 airfoil were made in the N.A.C.A 8-foot high-speed wind tunnel for the purpose of aiding in the development of a thick wing for high-speed airplanes. The tests were made at a lift coefficient of 0.1 for Reynolds Numbers from 1,750,000 to 8,690,000, corresponding to speeds from 80 to 440 miles per hour at 59 F. The effect on the profile drag of fixing the transition point was also investigated. The effect of compressibility on the rate of increase of pressure coefficients was found to be greater than that predicted by a simplified theoretical expression for thin wings. The results indicated that, for a lift coefficient of 0.1, the critical speed of the N.A.C,A. 45-125 airfoil was about 460 miles per hour at 59 F,. The value of the profile-drag coefficient at a Reynolds Number of 4,500,000 was 0.0058, or about half as large as the value for the N.A,C,A. 0025 airfoil. The increase in the profile-drag coefficient for a given movement of the transition point was about three times as large as the corresponding increase for the N.A.C,A. 0012 airfoil. Transition determinations indicated that, for Reynolds Numbers up to ?,000,000, laminar boundary 1ayers were maintained over approximately 40 percent of the upper and the lower surfaces of the airfoil.
Cooling system for a gas turbine using a cylindrical insert having V-shaped notch weirs
Grondahl, Clayton M.; Germain, Malcolm R.
1981-01-01
An improved cooling system for a gas turbine is disclosed. A plurality of V-shaped notch weirs are utilized to meter a coolant liquid from a pool of coolant into a plurality of platform and airfoil coolant channels formed in the buckets of the turbine. The V-shaped notch weirs are formed in a separately machined cylindrical insert and serve to desensitize the flow of coolant into the individual platform and airfoil coolant channels to design tolerances and non-uniform flow distribution.
NASA Astrophysics Data System (ADS)
Manela, A.; Halachmi, M.
2015-06-01
The acoustic signature of side-by-side airfoils, subject to small-amplitude harmonic pitching and incoming flow unsteadiness, is investigated. The two-dimensional near-field problem is formulated using thin-airfoil theory, where flow unsteadiness is modeled as a passing line vortex, and wake evolution is calculated via the Brown and Michael formula. Assuming that the setup is acoustically compact, acoustic radiation is obtained by means of the Powell-Howe acoustic analogy. The associated compact Green's function is calculated numerically using potential-flow analysis of the fluid-structure flow domain. Results, comparing the acoustic radiation of the double-airfoil system to a reference case of a single airfoil, point to several mechanisms of sound attenuation and sound amplification, caused by airfoil-airfoil and airfoils-wake interactions. It is found that counter-phase pitching of the airfoils results in effective cloaking of the system, which otherwise becomes significantly noisy (as a 5/2-power of the pitching frequency) at large frequencies. In addition, depending on the distance between airfoils, in-phase pitching may result in an acoustic signature equivalent to a single airfoil (when the airfoils are adjacent) or to two separate airfoils (when the airfoils are far apart). In general, flow unsteadiness produces more sound when interacting with a double (compared with a single) airfoil setup. However, airfoils' nonlinear wake-wake interactions give rise to a sound reduction mechanism, which becomes most efficient at times when incoming vorticity passes above airfoils' leading and trailing edges. The present scheme can be readily extended to consider the acoustic properties of various double-airfoil configurations, as well as multiple (> 2) airfoil setups.
A new parameterization of spectral and broadband ocean surface albedo.
Jin, Zhonghai; Qiao, Yanli; Wang, Yingjian; Fang, Yonghua; Yi, Weining
2011-12-19
A simple yet accurate parameterization of spectral and broadband ocean surface albedo has been developed. To facilitate the parameterization and its applications, the albedo is parameterized for the direct and diffuse incident radiation separately, and then each of them is further divided into two components: the contributions from surface and water, respectively. The four albedo components are independent of each other, hence, altering one will not affect the others. Such a designed parameterization scheme is flexible for any future update. Users can simply replace any of the adopted empirical formulations (e.g., the relationship between foam reflectance and wind speed) as desired without a need to change the parameterization scheme. The parameterization is validated by in situ measurements and can be easily implemented into a climate or radiative transfer model.
NASA Technical Reports Server (NTRS)
Hultgren, Lennart S.; Ashpis, David E.
2003-01-01
transition. Gad-el-Hak provides a review of various techniques for flow control in general and Volino discusses recent studies on separation control under low-pressure-turbine conditions utilizing passive as well as active devices. As pointed out by Volino, passive devices optimized for separation control at low Reynolds numbers tend to increase losses at high Reynolds numbers, Active devices have the attractive feature that they can be utilized only in operational regimes where they are needed and when turned off would not affect the flow. The focus in the present paper is an experimental Separation is induced on a flat plate installed in a closed-circuit wind tunnel by a shaped insert on the opposite wall. The flow conditions represent flow over the suction surface of a modem low-pressure-turbine airfoil ('Pak-B'). The Reynolds number, based on wetted plate length and nominal exit velocity, is varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low (0.2%) and high (2.5%) Gee-stream turbulence intensities are set using passive grids. A spanwise-oriented phased-plasma-array actuator, fabricated on a printed circuit board, is surface- flush-mounted upstream of the separation point and can provide forcing in a wide frequency range. Static surface pressure measurements and hot-wire anemometry of the base and controlled flows are performed and indicate that the glow-discharge plasma actuator is an effective device for separation control. of active separation control using glow discharge plasma actuators.
NASA Technical Reports Server (NTRS)
Maresh, J. L.; Bragg, M. B.
1984-01-01
A method has been developed to predict the contamination of an airfoil by insects and the resultant performance penalty. Insect aerodynamics have been modeled and the impingement of insects on an airfoil are solved by calculating their trajectories. Upon impact, insect rupture and the resulting height of the debris is determined based on experimental data. A boundary layer analysis is performed to determine which insects cause boundary layer transition and the resultant drag penalty. A contaminated airfoil figure of merit is presented to be used to compare airfoil susceptibility. Results show that the insect contamination effects depend on accretion conditions, airfoil angle of attack and Reynolds number. The importance of the stagnation region to designing airfoils for minimum drag penalties is discussed.
Influence of airfoil thickness on sound generated by high-frequency gust interactions
NASA Technical Reports Server (NTRS)
Tsai, C. T.; Kerschen, E. J.
1992-01-01
The sound radiated by interaction of a short wavelength gust with a symmetric thin airfoil is analyzed. The theory is based on a linearization of the Euler equations about the subsonic mean flow past the airfoil. The sound generation mechanism is found to be concentrated in a local region surrounding the parabolic nose of the airfoil; the size of this local region scales on the gust wavelength. At low Mach numbers, moderate values of airfoil thickness decrease the sound power, while at higher Mach numbers the sound power tends to increase with airfoil thickness. Airfoil thickness produces dramatic changes in the far field directivity. Both the sound power and the directivity are strong functions of the gust orientation.
Low-speed wind-tunnel results for symmetrical NASA LS(1)-0013 airfoil
NASA Technical Reports Server (NTRS)
Ferris, James C.; Mcghee, Robert J.; Barnwell, Richard W.
1987-01-01
A wind-tunnel test has been conducted in the Langley Low-Turbulence Pressure Tunnel to evaluate the performance of a symmetrical NASA LS(1)-0013 airfoil which is a 13-percent-thick, low-speed airfoil. The airfoil contour was obtained from the thickness distribution of a 13-percent-thick, high-performance airfoil developed for general aviation airplanes. The tests were conducted at Mach numbers from 0.10 tp 0.37 over a Reynolds number range from about 0.6 to 12.0 X 10 to the 6th power. The angle of attack varied from about -8 to 20 degrees. The results indicate that the aerodynamic characteristics of the present airfoil are similar to, but slightly better than, those of the NACA 0012 airfoil.
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.
Extensions and applications of a second-order landsurface parameterization
NASA Technical Reports Server (NTRS)
Andreou, S. A.; Eagleson, P. S.
1983-01-01
Extensions and applications of a second order land surface parameterization, proposed by Andreou and Eagleson are developed. Procedures for evaluating the near surface storage depth used in one cell land surface parameterizations are suggested and tested by using the model. Sensitivity analysis to the key soil parameters is performed. A case study involving comparison with an "exact" numerical model and another simplified parameterization, under very dry climatic conditions and for two different soil types, is also incorporated.
NASA Astrophysics Data System (ADS)
Pedersen, N. L.
2015-06-01
The strength of a gear is typically defined relative to durability (pitting) and load capacity (tooth-breakage). Tooth-breakage is controlled by the root shape and this gear part can be designed because there is no contact between gear pairs here. The shape of gears is generally defined by different standards, with the ISO standard probably being the most common one. Gears are manufactured using two principally different tools: rack tools and gear tools. In this work, the bending stress of involute teeth is minimized by shape optimization made directly on the final gear. This optimized shape is then used to find the cutting tool (the gear envelope) that can create this optimized gear shape. A simple but sufficiently flexible root parameterization is applied and emphasis is put on the importance of separating the shape parameterization from the finite element analysis of stresses. Large improvements in the stress level are found.
NASA Technical Reports Server (NTRS)
Nicks, Oran W.; Korkan, Kenneth D.
1991-01-01
Two reports on student activities to determine the properties of a new laminar airfoil which were delivered at a conference on soaring technology are presented. The papers discuss a wind tunnel investigation and analysis of the SM701 airfoil and verification of the SM701 airfoil aerodynamic charcteristics utilizing theoretical techniques. The papers are based on a combination of analytical design, hands-on model fabrication, wind tunnel calibration and testing, data acquisition and analysis, and comparison of test results and theory.
Wind Noise Measurements of Microphones Embedded in the Airfoil of an UAV
2012-10-01
Embedded in the Airfoil of an UAV October 2012 Wayne E. Prather and William G. Frazier National Center for Physical Acoustics...platforms has been increasing in recent years. Embedding microphones in the skin of an airfoil is one option for how acoustic sensors might be...employed on a small Unmanned Aerial Vehicle (UAV). The analysis of wind noise measurements made by an array of microphones embedded in the airfoil of a
A critical evaluation of the predictions of the NASA-Lockheed multielement airfoil computer program
NASA Technical Reports Server (NTRS)
Brune, G. W.; Manke, J. W.
1978-01-01
Theoretical predictions of several versions of the multielement airfoil computer program are evaluated. The computed results are compared with experimental high lift data of general aviation airfoils with a single trailing edge flap, and of airfoils with a leading edge flap and double slotted trailing edge flaps. Theoretical and experimental data include lift, pitching moment, profile drag and surface pressure distributions, boundary layer integral parameters, skin friction coefficients, and velocity profiles.
2011-05-13
sound production from a hydrofoil and identified three mechanisms: (1) low frequency curvature noise associated with interaction of a turbulent...2002). 2 Technical Approach A two-dimensional, dual-slotted, elliptic circulation control airfoil based on the hydrofoil studied by Rogers...airfoil, shown in Figure 1A, is designed based on the geometry of the hydrofoil previously studied by Rogers & Donnelly (2004). The airfoil’s profile
S830, S831, and S832 Airfoils: November 2001-November 2002
Somers, D. M.
2005-08-01
A family of quiet, thick, natural-laminar-flow airfoils, the S830, S831, and S832, for 40 - 50-meter-diameter, variable-speed/variable-pitch, horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, relatively insensitive to roughness, and low profile drag have been achieved. The airfoils should exhibit docile stalls, which meet the design goal. The constraints on the pitching moment and the airfoils thicknesses have been satisfied.
S833, S834, and S835 Airfoils: November 2001--November 2002
Somers, D. M.
2005-08-01
A family of quiet, thick, natural-laminar-flow airfoils, the S833, S834, and S835, for 1 - 3-meter-diameter, variable-speed/variable-pitch, horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, relatively insensitive to roughness, and low profile drag have been achieved. The airfoils should exhibit docile stalls, which meet the design goal. The constraints on the pitching moment and the airfoils thicknesses have been satisfied.
Systematic Airfoil Tests in the Large Wind Tunnel of the DVL
NASA Technical Reports Server (NTRS)
Doetsch, H; Kramer, M
1938-01-01
The present report is a description of systematic tests at maximum lift on airfoils with and without split flap and of profile drag at low lift. In order to obtain an opinion as to the suitability of the airfoils with flaps, the maximum-lift measurements were repeated on airfoils with split flaps. The profile drag at low lift was arrived at by direct weighing and momentum measurements and, since the profiles were of unusual depth, extended to large Reynolds numbers.
NASA Technical Reports Server (NTRS)
Mcalister, K. W.; Pucci, S. L.; Mccroskey, W. J.; Carr, L. W.
1982-01-01
Experimentally derived force and moment data are presented for eight airfoil sections that were tested at fixed and varying incidence in a subsonic two dimensional stream. Airfoil incidence was varied through sinusoidal oscillations in pitch over a wide range of amplitude and frequency. The surface pressure distribution, as well as the lift, drag, and pitching moment derived therefrom, are displayed in a uniform fashion to delineate the static and dynamic characteristics of each airfoil both in and out of stall.
Transition and Turbulence Structure in the Boundary Layers of an Oscillating Airfoil
1989-12-31
measurements near the trailing edge of an oscillating NACA 64A010 airfoil operating at various fiequencies and in unstalled condition. It was...to identify: (i) the conditions under which a sinusoidally oscillating NACA 0012 airfoil operates with a leading edge separation bubble; (ii) the...vortex formation. These experiments were conducted using an NACA 0012 airfoil with a tripping wire to promote transition of the boundary layer. The
Satellite-Based Model Parameterization of Diabatic Heating
NASA Technical Reports Server (NTRS)
Pielke, Roger, Sr.; Stokowski, David; Wang, Jih-Wang; Vukicevic, Tomislava; Leoncini, Giovanni; Matsui, Toshihisa; Castro, Christopher L.; Niyogi, Dev; Kishtawal, Chandra M.; Biazar, Arastoo; Doty, Kevin; McNider, Richard T.; Nair, Udaysankar; Tao, Wei-Kuo
2007-01-01
Future meteorological satellites are expected to provide much needed fine-scale information that can improve the accuracy of weather and climate models. As one application of this improved capability, we introduce the concept of a generalized parameterization framework using satellite datasets that will increase the accuracy and the computational efficiency of weather and climate modeling. In an atmospheric model, several different parameterizations usually are used to reproduce the various physical processes. However, it is generally unrealistic to separate the processes in this artificial way since the observations and physics make no such artificial separation. Thus, we propose a new unified parameterization framework to remove the unrealistic separation between parameterizations.
Airfoil for a turbine of a gas turbine engine
Liang, George
2010-12-21
An airfoil for a turbine of a gas turbine engine is provided. The airfoil comprises a main body comprising a wall structure defining an inner cavity adapted to receive a cooling air. The wall structure includes a first diffusion region and at least one first metering opening extending from the inner cavity to the first diffusion region. The wall structure further comprises at least one cooling circuit comprising a second diffusion region and at least one second metering opening extending from the first diffusion region to the second diffusion region. The at least one cooling circuit may further comprise at least one third metering opening, at least one third diffusion region and a fourth diffusion region.
Effect of oscillation frequency on wind turbine airfoil dynamic stall
NASA Astrophysics Data System (ADS)
Zhou, Z.; Li, C.; Nie, J. B.; Chen, Y.
2013-12-01
At the same oscillation amplitude, Reynolds Number, mean angle of attack, the dynamic stall characteristics of the NREL S809 airfoil undergoing sinusoidal pitch oscillations of different oscillation frequencies were investigated with modified k-ω SST turbulence model of CFD solution for two-dimensional numerical simulation. The predicted lift, drag coefficients and moment coefficients were compared with the Ohio State University wind tunnel test results, which showed a good agreement. The birth, development and breaking off of eddies were analyzed through streamline distribution around airfoil and the influence of oscillation frequencies on dynamic stall characteristics was also described and analyzed in detail, which enrich the database of dynamic stall characteristics needed by the quantization of oscillation frequencies on dynamic characteristics and prove that sliding mesh method is reliable when dealing with dynamic stall problems.
Kasprzyk airfoil. The first wind-tunnel tests
NASA Technical Reports Server (NTRS)
Wusatowski, T.
1984-01-01
The Kasprzyk slotted flap glider airfoil (the Kasper wing) enabling glider flight at 32 km/h and 0.5 m/sec descent speed was wind tunnel tested in the U.S. The test layout is described and reasons offered for discrepancies between wind tunnel results and Polish in flight data: high induced drag caused by relative size of model wing span and tunnel, by vortex attenuators on the model and their proximity to the tunnel wall, nonsimilarity between flow over a smooth wing and flow over the Kasprzyk wing with bound vortices, obstruction of the tunnel test chamber cross section by the model wing, discrepant Reynolds numbers, and model airfoil aspect ratio much smaller than the prototype. The overall results offer partial confirmation of the Kasprzyk theory, but further in tunnel and in flight studies are recommended.
CAST-10-2/DOA 2 Airfoil Studies Workshop Results
NASA Technical Reports Server (NTRS)
Ray, Edward J. (Compiler); Hill, Acquilla S. (Compiler)
1989-01-01
During the period of September 23 through 27, 1988, the Transonic Aerodynamics Division at the Langely Research Center hosted an International Workshop on CAST-10-2/DOA 2 Airfoil Studies. The CAST-10 studies were the outgrowth of several cooperative study agreements among the NASA, the NAE of Canada, the DLR of West Germany, and the ONERA of France. Both theoretical and experimental CAST-10 airfoil results that were obtained form an extensive series of tests and studies, were reviewed. These results provided an opportunity to make direct comparisons of adaptive wall test section (AWTS) results from the NASA 0.3-meter Transonic Cryogenic Tunnel and ONERA T-2 AWTS facilities with conventional ventilated wall wind tunnel results from the Canadian high Reynolds number two-dimensional test facility. Individual papers presented during the workshop are included.
Theory and Low-Order Modeling of Unsteady Airfoil Flows
NASA Astrophysics Data System (ADS)
Ramesh, Kiran
Unsteady flow phenomena are prevalent in a wide range of problems in nature and engineering. These include, but are not limited to, aerodynamics of insect flight, dynamic stall in rotorcraft and wind turbines, leading-edge vortices in delta wings, micro-air vehicle (MAV) design, gust handling and flow control. The most significant characteristics of unsteady flows are rapid changes in the circulation of the airfoil, apparent-mass effects, flow separation and the leading-edge vortex (LEV) phenomenon. Although experimental techniques and computational fluid dynamics (CFD) methods have enabled the detailed study of unsteady flows and their underlying features, a reliable and inexpensive loworder method for fast prediction and for use in control and design is still required. In this research, a low-order methodology based on physical principles rather than empirical fitting is proposed. The objective of such an approach is to enable insights into unsteady phenomena while developing approaches to model them. The basis of the low-order model developed here is unsteady thin-airfoil theory. A time-stepping approach is used to solve for the vorticity on an airfoil camberline, allowing for large amplitudes and nonplanar wakes. On comparing lift coefficients from this method against data from CFD and experiments for some unsteady test cases, it is seen that the method predicts well so long as LEV formation does not occur and flow over the airfoil is attached. The formation of leading-edge vortices (LEVs) in unsteady flows is initiated by flow separation and the formation of a shear layer at the airfoil's leading edge. This phenomenon has been observed to have both detrimental (dynamic stall in helicopters) and beneficial (high-lift flight in insects) effects. To predict the formation of LEVs in unsteady flows, a Leading Edge Suction Parameter (LESP) is proposed. This parameter is calculated from inviscid theory and is a measure of the suction at the airfoil's leading edge. It
Supersonic flow past pointed-nose thin airfoils
NASA Astrophysics Data System (ADS)
Hemdan, Hamdi T.; Selim, Abdulfatah A.
1992-05-01
Accurate results for steady inviscid supersonic flow past any pointed-nose thin airfoil at small angle of attack are obtained in this paper. Straightforward perturbation from some basic wedge flow is applied to a recently developed hypersonic theory with second-order terms in the longitudinal curvature included. The hypersonic equations are thus reduced to two systems of simple linear ordinary differential equations for which polynomial solutions of the first and second degree are obtained. Further, the hypersonic theory is extended to include second-order effects on the basic wedge flow. This resulted in great increase in the accuracy of the results for curved airfoils besides, the range of validity of the theory extended to cover moderate supersonic speeds with quite reasonable accuracy.
Horizontal axis wind turbine post stall airfoil characteristics synthesization
NASA Technical Reports Server (NTRS)
Tangler, James L.; Ostowari, Cyrus
1995-01-01
Blade-element/momentum performance prediction codes are routinely used for wind turbine design and analysis. A weakness of these codes is their inability to consistently predict peak power upon which the machine structural design and cost are strongly dependent. The purpose of this study was to compare post-stall airfoil characteristics synthesization theory to a systematically acquired wind tunnel data set in which the effects of aspect ratio, airfoil thickness, and Reynolds number were investigated. The results of this comparison identified discrepancies between current theory and the wind tunnel data which could not be resolved. Other factors not previously investigated may account for these discrepancies and have a significant effect on peak power prediction.
Predicting aerodynamic characteristic of typical wind turbine airfoils using CFD
Wolfe, W.P.; Ochs, S.S.
1997-09-01
An investigation was conducted into the capabilities and accuracy of a representative computational fluid dynamics code to predict the flow field and aerodynamic characteristics of typical wind-turbine airfoils. Comparisons of the computed pressure and aerodynamic coefficients were made with wind tunnel data. This work highlights two areas in CFD that require further investigation and development in order to enable accurate numerical simulations of flow about current generation wind-turbine airfoils: transition prediction and turbulence modeling. The results show that the laminar-to turbulent transition point must be modeled correctly to get accurate simulations for attached flow. Calculations also show that the standard turbulence model used in most commercial CFD codes, the k-e model, is not appropriate at angles of attack with flow separation. 14 refs., 28 figs., 4 tabs.