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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Wind tunnel studies of circulation control elliptical airfoils

Effects of blown jets on the lift and drag of cambered elliptical airfoils are described. Performance changes due to a splitter plate attached to the lower surface of an elliptical airfoil near the trailing edge with and without blowing are indicated. Lift and drag characteristics of airfoils with two blown jets are compared with airfoils with single blowing jets. Airfoil designs that vary the location of a second jet relative to a fixed jet are described.

Franke, M. E.↗
NASA NTRS1988

The NASA Langley Laminar-Flow-Control (LFC) experiment on a swept, supercritical airfoil: Design overview

A large-chord, swept, supercritical, laminar-flow-control (LFC) airfoil was designed and constructed and is currently undergoing tests in the Langley 8 ft Transonic Pressure Tunnel. The experiment was directed toward evaluating the compatibility of LFC and supercritical airfoils, validating prediction techniques, and generating a data base for future transport airfoil design as part of NASA's ongoing research program to significantly reduce drag and increase aircraft efficiency. Unique features of the airfoil included a high design Mach number with shock free flow and boundary layer control by suction. Special requirements for the experiment included modifications to the wind tunnel to achieve the necessary flow quality and contouring of the test section walls to simulate free air flow about a swept model at transonic speeds. Design of the airfoil with a slotted suction surface, the suction system, and modifications to the tunnel to meet test requirements are discussed.

Harris, Charles D.↗
NASA NTRS1986

High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

Johnson, Charles B.↗
NASA NTRS1989

Supercritical flow past a symmetrical bicircular arc airfoil

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.

Holt, Maurice↗
NASA NTRS1989

Numerical solution of periodic vortical flows about a thin airfoil

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.

Scott, James R.↗
NASA NTRS1989

Prop-fan airfoil icing characteristics

An icing test program was carried out during the 1986-87 and 1987-88 icing seasons at the Fluidyne Icing Tunnel. The testing consisted of evaluating the icing characteristics and aerodynamic performance characteristics of two thin, two-dimensional airfoil sections. The two airfoils were a NACA Series 16-203 and a Hamilton Standard PF1-304, 3 and 4 percent thickness-to-chord ratio airfoils, respectively. These airfoils are representative of the thin airfoils currently being evaluated for Prop-Fan propulsion systems. The tests were conducted over a wide range of icing conditions, angles-of-attack, and Mach numbers (0.3 to 0.8).

Pike, J. A.↗
NASA NTRS1989

Optimization of natural laminar flow airfoils for high section lift-to-drag ratios in the lower Reynolds number range

Relatively thin natural-laminar-flow airfoils were arranged optimally for different design lift coefficients in the wing chord Reynolds number ranges of 200,000-600,00 and 0.875 x 10 to the 6th to 2 x 10 to the 6th. The 9.5 percent thick airfoil ASM-LRN-010, the 7.9 percent thick airfoil ASM-LRN-012, the 10.4 percent thick airfoil ASM-LRN-015, and the 8.2 percent thick airfoil ASM-LRN-017 were designed for high lift-to-drag ratios using Drela's design and analysis.

Pfenninger, Werner↗
NASA NTRS1989

Numerical solution of periodic vortical flows about a thin airfoil

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 Sommerfield radiation condition is applied to the unsteady pressure on the grid line corresponding to the far field boundary. The results are compared with a Possion 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.

Scott, James R.↗
NASA NTRS1989

Experience with some repeat tests on the 9 inch chord CAST-10-2/DOA 2 airfoil model in the Langley 0.3-m TCT adaptive wall test section

A co-operative testing program is in progress between the Langley Research Center (NASA) and the National Aeronautical Establishment (NAE, Canada) to validate two different techniques of airfoil testing at transonic speeds. The procedure employed is to test the same airfoil model in the NAE two-dimensional tunnel and the Langley 0.3-m Transonic Cryogenic Tunnel (0.3-m TCT). The airfoil model used in testing was CAST-10-2/DOA-2 super-critical airfoil. The Langley 0.3-m TCT has a relatively small cross section of 13 in x 13 in, giving a (h/c) ratio of 1.44 for the same 9 in chord model. The approach employed in the 0.3-m TCT aims towards eliminating the wall effects by using active walls. The top and bottom walls are flexible. By changing the wall shapes during a test in an iterative manner, the wall interference effects are reduced. The method employed to change the wall shapes is the adaptive wall technique. The current test program provided an opportunity to validate the adaptive wall technique in the 0.3-m TCT. The relatively long chord airfoil represents a severe test case to test the efficacy of the adaptive wall technique under cryogenic conditions. The program also involved removal of side wall boundary-layer thus increasing the complexity of the wall adaptation technique. This paper deals with some salient results obtained regarding repeatability of test data and possible residual interference effects.

Murthy, A. V.↗
NASA NTRS1990

Numerical computation of viscous flow about unconventional airfoil shapes

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.

Ahmed, S.↗
NASA NTRS1991

Wall-interference assessment and corrections for transonic NACA 0012 airfoil data from various wind tunnels

A nonlinear, four wall, post-test wall interference assessment/correction (WIAC) code was developed for transonic airfoil data from solid wall wind tunnels with flexibly adaptable top and bottom walls. The WIAC code was applied over a broad range of test conditions to four sets of NACA 0012 airfoil data, from two different adaptive wall wind tunnels. The data include many test points for fully adapted walls, as well as numerous partially adapted and unadapted test points, which together represent many different model/tunnel configurations and possible wall interference effects. Small corrections to the measured Mach numbers and angles of attack were obtained from the WIAC code even for fully adapted data; these corrections generally improve the correlation among the various sets of airfoil data and simultaneously improve the correlation of the data with calculations for a 2-D, free air, Navier-Stokes code. The WIAC corrections for airfoil data taken in fully adapted wall test sections are shown to be significantly smaller than those for comparable airfoil data from straight, slotted wall test sections. This indicates, as expected, a lesser degree of wall interference in the adapted wall tunnels relative to the slotted wall tunnels. Application of the WIAC code to this data was, however, somewhat more difficult and time consuming than initially expected from similar previous experience with WIAC applications to slotted wall data.

Green, Lawrence L.↗
NASA NTRS1990

Influence of airfoil nose radius on sound generated by gust interactions

The sound radiated by interaction of a short wavelength gust (k much greater than 1) 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. Airfoil thickness produces terms of relative order O(sq rt S) in the far field sound, where S is a Strouhal number based on the airfoil nose radius. At low Mach numbers, entropy gusts generate much less sound than vorticity gusts. However, for O(1) Mach numbers and moderate values of S, entropy and vorticity gusts produce similar sound levels.

Tsai, C. T.↗
NASA NTRS1991

A multi-point inverse airfoil design method based on conformal mapping

An exact method of multipoint inverse airfoil design for incompressible flow is presented. Multipoint design is handled by dividing the airfoil into a number of desired segments. For each segment the velocity distribution is prescribed together with an angle of attack at which the prescribed velocity distribution is to be achieved. In this manner, multipoint design objectives can be taken into account in the initial specification of the velocity distribution. In order for the multipoint inverse airfoil design problem to be well posed, three integral constraints and several conditions arise which must be satisfied. Further restrictions are imposed if the airfoil is to have a specified pitching moment, thickness ratio, or other constraints. The system of equations is solved partly as a linear system and partly through multidimensional Newton iteration. Since the velocity distribution is prescribed about the circle in the angular coordinate, specification of the velocity in terms of arc length is handled through the multidimensional Newton iteration as well. The current formulation sets the stage for a more general multipoint inverse airfoil design method in which it will be possible to specify the velocity distribution, some boundary-layer development, or the surface geometry along a segment.

Selig, Michael S.↗
NASA NTRS1990

An airfoil design method for viscous flows

An airfoil design procedure is described that has been incorporated into an existing two-dimensional Navier-Stokes airfoil analysis method. The resulting design method, an iterative procedure based on a residual-correction algorithm, permits the automated design of airfoil sections with prescribed surface pressure distributions. This paper describes the inverse design method and the technique used to specify target pressure distributions. An example airfoil design problem is described to demonstrate application of the inverse design procedure. It shows that this inverse design method develops useful airfoil configurations with a reasonable expenditure of computer resources.

Malone, J. B.↗
NASA NTRS1991

Euler study on porous transonic airfoils with a view toward multipoint design

Euler solutions for steady transonic flow (Free-stream Mach 0.63-0.8, alpha = 0-2 deg) over NACA 0012 and supercritical airfoils with solid as well as porous surfaces suggest porosity as a means to realize multipoint design for transonic airfoils. The porous surfaces extend over at least 90 percent of the chord. The porosity distribution is described by a modified sine wave with several amplitudes. Either connected or separated cavities are assumed to lie underneath the upper and lower surfaces. Applied to an NACA 0012 airfoil, porosity generally increases lift, in some instances by up to 65 percent. Porous NACA 0012 airfoils in supercritical flow yield reductions of an order of magnitude in wave drag at constant lift, compared to their solid counterpart. Making the surface of a supercritical airfoil permeable also leads to sizeable reductions in wave drag at constant lift for overspeed conditions. The discussion of the computed results addresses issues such as grid sensitivity and checks for systematic errors.

Hartwich, Peter M.↗
NASA NTRS1992

Efficient simulation of incompressible viscous flow over single and multi-element airfoils

Incompressible viscous turbulent flows over single- and multiple-element airfoils are numerically simulated in an efficient manner by solving the incompressible Navier-Stokes equations. The solution algorithm uses the method of pseudocompressibility with an upwind-differencing scheme for the convective fluxes and an implicit line-relaxation scheme to study high-lift take-off and landing configurations and to compute lift and drag at various angles of attack up to stall. Two different turbulence models are tested in computing the flow over an NACA 4412 airfoil. The approach used for multiple-element airfoils involves the use of multiple zones of structured grids fitted to each element. Two different approaches are compared: a patched system of grids and an overlaid Chimera system of grids. Computational results are presented for two-element, three-element, and four-element airfoil configurations. Excellent agreement with experimental surface-pressure coefficients is seen. The code converges in less than 200 iterations, requiring on the order of one minute of CPU time on a CRAY YMP per element in the airfoil configuration.

Rogers, Stuart E.↗
NASA NTRS1992

Recent progress in the analysis of iced airfoils and wings

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.

Cebeci, Tuncer↗
NASA NTRS1992

Efficient simulation of incompressible viscous flow over multi-element airfoils

The incompressible, viscous, turbulent flow over single and multi-element airfoils is numerically simulated in an efficient manner by solving the incompressible Navier-Stokes equations. The computer code uses the method of pseudo-compressibility with an upwind-differencing scheme for the convective fluxes and an implicit line-relaxation solution algorithm. The motivation for this work includes interest in studying the high-lift take-off and landing configurations of various aircraft. In particular, accurate computation of lift and drag at various angles of attack, up to stall, is desired. Two different turbulence models are tested in computing the flow over an NACA 4412 airfoil; an accurate prediction of stall is obtained. The approach used for multi-element airfoils involves the use of multiple zones of structured grids fitted to each element. Two different approaches are compared: a patched system of grids, and an overlaid Chimera system of grids. Computational results are presented for two-element, three-element, and four-element airfoil configurations. Excellent agreement with experimental surface pressure coefficients is seen. The code converges in less than 200 iterations, requiring on the order of one minute of CPU time (on a CRAY YMP) per element in the airfoil configuration.

Rogers, Stuart E.↗