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At least 1,261 records · Page 70

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NASA NTRS1994

Control theory based airfoil design using the Euler equations

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.

Jameson, Antony↗

NASA NTRS1994

Lift enhancing tabs for airfoils

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.

Ross, James C.↗

NASA NTRS1995

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

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

Allison, Dennis O.↗

NASA NTRS1995

Horizontal axis wind turbine post stall airfoil characteristics synthesization

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.

Tangler, James L.↗

NASA NTRS1995

Numerical study to assess sulfur hexafluoride as a medium for testing multielement airfoils

A methodology is described for computing viscous flows of air and sulfur hexafluoride (SF6). The basis is an existing flow solver that calculates turbulent flows in two dimensions on unstructured triangular meshes. The solver has been modified to incorporate the thermodynamic model for SF6 and used to calculate the viscous flow over two multielement airfoils that have been tested in a wind tunnel with air as the test medium. Flows of both air and SF6 at a free-stream Mach number of 0.2 and a Reynolds number of 9 x 10(exp 6) are computed for a range of angles of attack corresponding to the wind-tunnel test. The computations are used to investigate the suitability of SF6 as a test medium in wind tunnels and are a follow-on to previous computations for single-element airfoils. Surface-pressure, lift, and drag coefficients are compared with experimental data. The effects of heavy gas on the details of the flow are investigated based on computed boundary-layer and skin-friction data. In general, the predictions in SF6 vary little from those in air. Within the limitations of the computational method, the results presented are sufficiently encouraging to warrant further experiments.

Bonhaus, Daryl L.↗

NASA NTRS1994

Effect of Reynolds number and turbulence on airfoil aerodynamics at -90-degree incidence

A method has been developed for calculating the viscous flow about airfoils with and without deflected flaps at -90 deg incidence. This method provides for the solution of the unsteady incompressible Navier-Stokes equations by means of an implicit technique. The solution is calculated on a body-fitted computational mesh using a staggered-grid method. The vorticity is defined at the node points, and the velocity components are defined at the mesh-cell sides. The staggered-grid orientation provides for accurate representation of vorticity at the node points and the continuity equation at the mesh-cell centers. The method provides for the noniterative solution of the flowfield and satisfies the continuity equation to machine zero at each time step. The method is evaluated in terms of its stability to predict two-dimensional flow about an airfoil at -90-deg incidence for varying Reynolds number and laminar/turbulent models. The variations of the average loading and surface pressure distribution due to flap deflection, Reynolds number, and laminar or turbulent flow are presented and compared with experimental results. The comparisom indicate that the calculated drag and drag reduction caused by flap deflection and the calculated average surface pressure are in excellent agreement with the measured results at a similar Reynolds number.

Stremel, Paul M.↗

NASA NTRS1994

Low-dimensional description of the dynamics in separated flow past thick airfoils

Results are presented for the numerical simulation of unsteady viscous incompressible flow past thick airfoils. Specifically, flow past a NACA 4424 at an angle of attack of 2.5 deg and Reynolds numbers in the range of 1700-4000 has been simulated using the spectral element method. At these conditions the flow is separatedd and an unsteady wake is formed. Application of the method of empirical eigenfunction reveals the structure of the most energetic components of the flow. These are found to occur in pairs that, through phase exchange, are responsible for the vortex shedding. A set of ordinary differential equations is obtained for the amplitudes of these eigenfunctions by a Galerkin projection of the Navier-Stokes equations. The solutions of the model system are compared with the full simulation. The work is of relevance to the transition process and observed routes to chaos in airfoil wakes.

Deane, Anil E.↗

NASA NTRS1992

Numerical study of multi-element airfoil aerodynamics

Unsteady flowfields around oscillating Boeing VR7 airfoil with and without a leading-edge slat were numerically investigated by a novel zonal method using a conformal mapping technique. Numerical aero-dynamic hysteresis loops show that the leading-edge slat prevents the airfoil dynamic stall at reduced frequency of 0.15, Reynolds number of 1 million, and the oscillation range of 5 deg to 25 deg.

Wang, Clin M.↗

NASA NTRS1995

Temperature Measurements Reveal Accretion Of Ice On Airfoils

Method of detecting accretion of ice on airfoil based on measurement of difference between temperatures at icing and nonicing locations on airfoil. Release of latent heat of fusion gives rise to characteristic gradient of temperature. Temperatures and difference between them measured by use of contact sensors or remote sensor. Measurements used to trigger alarms and/or deicing equipment.

Hansman, Robert J., Jr.↗

NASA NTRS1995

A comparison of design variables for control theory based airfoil optimization

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.

Reuther, James↗

NASA NTRS1995

Tables of properties of airfoil polynomials

This monograph provides an extensive list of formulas for airfoil polynomials. These polynomials provide convenient expansion functions for the description of the downwash and pressure distributions of linear theory for airfoils in both steady and unsteady subsonic flow.

Desmarais, Robert N.↗

NASA NTRS1996

A Hybrid Boundary Element-Finite Volume Method for Unsteady Transonic Airfoil Flows

A hybrid boundary element finite volume method for unsteady transonic flow computation has been developed. In this method, the unsteady Euler equations in a moving frame of reference are solved in a small embedded domain (inner domain) around the airfoil using an implicit finite volume scheme. The unsteady full-potential equation, written in the same frame of reference and in the form of the Poisson equation. is solved in the outer domain using the integral equation boundary element method to provide the boundary conditions for the inner Euler domain. The solution procedure is a time-accurate stepping procedure, where the outer boundary conditions for the inner domain are updated using the integral equation -- boundary element solution over the outer domain. The method is applied to unsteady transonic flows around the NACA0012 airfoil undergoing pitching oscillation and ramp motion. The results are compared with those of an implicit Euler equation solver, which is used throughout a large computational domain, and experimental data.

Hu, Hong↗

NASA NTRS1995

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

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

Kaltenbach, Hans-Jakob↗

NASA NTRS1996

The modelling of symmetric airfoil vortex generators

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.

Reichert, B. A.↗

NASA NTRS1996

Improving CAP-TSD steady pressure solutions through airfoil slope modification

A modification of airfoil section geometry is examined for improvement of the leading edge pressures predicted by the Computational Aeroelasticity Program - Transonic Small Disturbance (CAP-TSD). Results are compared with Eppler solutions to assess improvement. Preliminary results indicate that a fading function modification of section slopes is capable of significant improvements in the pressures near the leading edge computed by CAP-TSD. Application of this modification to airfoil geometry before use in CAP-TSD is shown to reduce the nonphysical pressure peak predicted by the transonic small disturbance solver. A second advantage of the slope modification is the substantial reduction in sensitivity of CAP-TSD steady pressure solutions to the computational mesh.

Mitterer, Kent F.↗

NASA NTRS1996

Performance Improvement Through Indexing of Turbine Airfoils: Numerical Simulation - Part 2

An experimental/analytical study has been conducted to determine the performance improvements achievable by circumferentially indexing succeeding rows of turbine stator airfoils. A series of tests was conducted to experimentally investigate stator wake clocking effects on the performance of the space shuttle main engine (SSME) alternate turbopump development (ATD) fuel turbine test article (TTA). The results from this study indicate that significant increases in stage efficiency can be attained through application of this airfoil clocking concept. Details of the experiment and its results are documented in part 1 of this paper. In order to gain insight into the mechanisms of the performance improvement, extensive computational fluid dynamics (CFD) simulations were executed. The subject of the present paper is the initial results from the CFD investigation of the configurations and conditions detailed in part 1 of the paper. To characterize the aerodynamic environments in the experimental test series, two-dimensional (2D), time accurate, multistage, viscous analyses were performed at the TTA midspan. Computational analyses for five different circumferential positions of the first stage stator have been completed. Details of the computational procedure and the results are presented. The analytical results verify the experimentally demonstrated performance improvement and are compared with data whenever possible. Predictions of time-averaged turbine efficiencies as well as gas conditions throughout the flow field are presented. An initial understanding of the turbine performance improvement mechanism based on the results from this investigation is described.

Griffin, Lisa W.↗

NASA NTRS1995

Analysis of a High-Lift Multi-Element Airfoil using a Navier-Stokes Code

A thin-layer Navier-Stokes code, CFL3D, was utilized to compute the flow over a high-lift multi-element airfoil. This study was conducted to improve the prediction of high-lift flowfields using various turbulence models and improved glidding techniques. An overset Chimera grid system is used to model the three element airfoil geometry. The effects of wind tunnel wall modeling, changes to the grid density and distribution, and embedded grids are discussed. Computed pressure and lift coefficients using Spalart-Allmaras, Baldwin-Barth, and Menter's kappa-omega - Shear Stress Transport (SST) turbulence models are compared with experimental data. The ability of CFL3D to predict the effects on lift coefficient due to changes in Reynolds number changes is also discussed.

Whitlock, Mark E.↗

NASA NTRS1997

PROFILE: Airfoil Geometry Manipulation and Display

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.

Collins, Leslie↗