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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 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

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

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

Ayers, T. G.↗
NASA NTRS1983

An analytical evaluation of the icing properties of several low and medium speed airfoils

A droplet trajectory computer code is used to analyze the water droplet impingement characteristics of several low and medium speed airfoils. The results are reported in terms of the airfoil maximum impingement efficiency, and limits of impingement. The airfoils are analyzed for angles of attack of -4 to 12 degrees, and modified inertia parameters of 0.01 to 1.0. These data are then used to evaluate the effect of airfoil geometry: leading edge radius, maximum thickness, maximum camber, and angle of attack on airfoil icing characteristics. Airfoils from the NACA four digit series to the new NASA NLF airfoils are considered. The strongest correlations were seen between airfoil leading edge radius and maximum local collection efficiency, thickness and total collection efficiency, and camber and the limits of impingement.

Bragg, M. B.↗
NASA NTRS1983

Flow fields and aerodynamic characteristics for hypersonic missiles with mid-fuselage inlets

A study was made to quantify forebody flow fields and to evaluate aerodynamic performance trends on a matrix of fuselage shapes for the mid-inlet/bolt-on-engine class of hypersonic airbreathing missiles for the Navy's vertical box launcher. The study indicated that inlet mass flow and pressure recovery can be increased by cambering the nose and increasing the width of the fuselage at both Mach 4 acceleration and Mach 6 cruise conditions. Aerodynamic trim predictions show that the drag at zero lift at Mach 4 decreases while the L/D max at Mach 6 increases with the nose camber, although these tendencies reverse with increasing width of maximum fuselage cross section.

Hunt, J. L.↗
NASA NTRS1984

Effect of geometry on airfoil icing characteristics

A droplet trajectory computer code is used to predict the water droplet impingement characteristics of several low- and medium-speed airfoils. The maximum impingement efficiency, total collection efficiency, and limits of impingement are analyzed as functions of the airfoil geometry and freestream conditions. The airfoil geometry is represented by leading edge radius, maximum thickness, maximum camber, and angle of attack. The analysis shows that the primary effects are an increase in maximum impingement efficiency with a decrease in leading edge radius, a reduction in total collection efficiency for thicker airfoils, and a change in the limits of impingement for airfoils of different maximum camber.

Bragg, M. B.↗
NASA NTRS1982

Experimental trim drag values and flow-field measurements for a wide-body transport model with conventional and supercritical wings

The purpose of this study was to determine if advanced supercritical wings incur higher trim drag values at cruise conditions than current wide body technology wings. Relative trim drag increments were measured in an experimental wind tunnel investigation conducted in the Langley 8 Foot Transonic Pressure Tunnel. The tests utilized a high aspect ratio supercritical wing and a wide body aircraft wing, in conjunction with five different horizontal tail configurations, mounted on a representative wide body fuselage. The three low tail and two T-tail configurations were designed to measure the effects of horizontal tail size, location, and camber on the trim drag increments for the two wings. Longitudinal force and moment data were taken at a Mach number of 0.82 and design cruise lift coefficients for the wide body and supercritical wings of 0.45 and 0.55, respectively. The data indicate that the supercritical wing does not have significantly higher trim drag than the wide body wing. A reduction in tail size, combined with relaxed static stability, produced trim drag reductions for both wings. The cambered tails had higher trim drag increments than the symmetrical tails for both wings, and the T-tail configurations had lower trim drag increments than the low tail configurations.

Jacobs, P. F.↗
NASA NTRS1982

VORCOR: A computer program for calculating characteristics of wings with edge vortex separation by using a vortex-filament and-core model

A computer code base on an improved vortex filament/vortex core method for predicting aerodynamic characteristics of slender wings with edge vortex separations is developed. The code is applicable to camber wings, straked wings or wings with leading edge vortex flaps at subsonic speeds. The prediction of lifting pressure distribution and the computer time are improved by using a pair of concentrated vortex cores above the wing surface. The main features of this computer program are: (1) arbitrary camber shape may be defined and an option for exactly defining leading edge flap geometry is also provided; (2) the side edge vortex system is incorporated.

Pao, J. L.↗
NASA NTRS1983

Computer programs for smoothing and scaling airfoil coordinates

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.

Morgan, H. L., Jr.↗
NASA NTRS1985

TRO-2D - A code for rational transonic aerodynamic optimization

Features and sample applications of the transonic rational optimization (TRO-2D) code are outlined. TRO-2D includes the airfoil analysis code FLO-36, the CONMIN optimization code and a rational approach to defining aero-function shapes for geometry modification. The program is part of an effort to develop an aerodynamically smart optimizer that will simplify and shorten the design process. The user has a selection of drag minimization and associated minimum lift, moment, and the pressure distribution, a choice among 14 resident aero-function shapes, and options on aerodynamic and geometric constraints. Design variables such as the angle of attack, leading edge radius and camber, shock strength and movement, supersonic pressure plateau control, etc., are discussed. The results of calculations of a reduced leading edge camber transonic airfoil and an airfoil with a natural laminar flow are provided, showing that only four design variables need be specified to obtain satisfactory results.

Davis, W. H., Jr.↗
NASA NTRS1987

Investigation of generic hub fairing and pylon shapes to reduce hub drag

Reported are investigations of fairing configurations pointed toward substantially reducing hub drag. Experimental investigations have shown the importance of hub-fairing camber, lower-surface curvature, and relative size of the drag. The significance of pylon and hub fairings in combination have also been shown. Model test data presented here documented these findings, and also showed the effect of gaps and hub-fairing inclination angle on drag. From a drag standpoint, the best hub fairing had a circular arc, upper-surface curvature, a flat bottom surface, and 8.75% camber.

Stroub, Robert H.↗
NASA NTRS1987

Reduction of hub- and pylon-fairing drag

A wind tunnel test was conducted to obtain data on several drag reduction methods for rotorcraft hubs. The objective of the test was to use small-scale models to develop the technology to substantially reduce hub drag. A helicopter test model, which did not incorporate a rotor, was used to study single-rotor fairing configurations. The rotor shaft assembly was modeled with nonrotating hardware. Drag trend data was obtained for the following fairing configuration changes: hub-fairing camber, hub-fairing thickness ratio, hub-fairing surface curvature, hub-fairing height with respect to the fuselage, inclusion of blade shanks in the hub fairings, hub- and pylon-fairing gaps, pylon-fairing cross-sectional geometry, pylon-fairing thickness ratio and camber. Substantial drag reductions could potentially result from application of the test results to helicopter hub- and pylon-fairing design.

Young, Larry A.↗
NASA NTRS1989

Helicopter hub fairing and pylon interference drag

A wind tunnel test was conducted to study the aerodynamics of helicopter hub and pylon fairings. The test was conducted in the 7-by 10 Foot Subsonic Wind Tunnel (Number 2) at Ames Research Center using a 1/5-scale XH-59A fuselage model. The primary focus of the test was on the rotor hub fairing and pylon mutual interference drag. Parametric studies of pylon and hub fairing geometry were also conducted. This report presents the major findings of the test as well as tabulated force and moment data, flow visualization photographs, and graphical presentations of the drag data. The test results indicate that substantial drag reduction can be attained through the use of a cambered hub fairing with circular arc upper surface and flat lower surface. Furthermore, a considerable portion of the overall drag reduction is attributed to the reduction in the hub-on-pylon interference drag. It is also observed that the lower surface curvature of the fairing has a strong influence on the hub fairing and on pylon interference drag. However, the drag reduction benefit that was obtained by using the cambered hub fairing with a flat lower surface was adversely affected by the clearance between the hub fairing and the pylon.

Graham, D. R.↗
NASA NTRS1987

Investigation of generic hub fairing and pylon shapes to reduce hub drag

This paper reports investigations of fairing configurations pointed toward substantially reducing hub drag. Experimental investigations have shown the importance of hub-fairing camber,lower-surface curvature, and relative size on the drag. The significance of pylon and hub fairings in combination have also been shown. Model test data presented here documented these findings, and also showed the effect of gaps and hub-fairing inclination angle on drag. From a drag standpoint, the best hub fairing had a circular arc, upper-surface curvature, a flat bottom surface, and 8.75 percent camber.

Stroub, Robert H.↗
NASA NTRS1989

An Experimental Investigation of Helicopter Rotor Hub Fairing Drag Characteristics

A study was done in the NASA 14- by 22-Foot Wind Tunnel at Langley Research Center on the parasite drag of different helicopter rotor hub fairings and pylons. Parametric studies of hub-fairing camber and diameter were conducted. The effect of hub fairing/pylon clearance on hub fairing/pylon mutual interference drag was examined in detail. Force and moment data are presented in tabular and graphical forms. The results indicate that hub fairings with a circular-arc upper surface and a flat lower surface yield maximum hub drag reduction; and clearance between the hub fairing and pylon induces high mutual-interference drag and diminishes the drag-reduction benefit obtained using a hub fairing with a flat lower surface. Test data show that symmetrical hub fairings with circular-arc surfaces generate 74 percent more interference drag than do cambered hub fairings with flat lower surfaces, at moderate negative angle of attack.

Sung, D. Y.↗
NASA NTRS1990

The effect of steady aerodynamic loading on the flutter stability of turbomachinery blading

An aeroelastic analysis is presented which accounts for the effect of steady aerodynamic loading on the aeroelastic stability of a cascade of compressor blades. The aeroelastic model is a two degree of freedom model having bending and torsional displacements. A linearized unsteady potential flow theory is used to determine the unsteady aerodynamic response coefficients for the aeroelastic analysis. The steady aerodynamic loading was caused by the addition of airfoil thickness and camber and steady flow incidence. The importance of steady loading on the airfoil unsteady pressure distribution is demonstrated. Additionally, the effect of steady loading on the tuned flutter behavior and flutter boundaries indicates that neglecting either airfoil thickness, camber or incidence could result in nonconservative estimates of flutter behavior.

Smith, Todd E.↗
NASA NTRS1991

Static thrust-vectoring performance of nonaxisymmetric convergent-divergent nozzles with post-exit yaw vanes

A static (wind-off) test was conducted in the Static Test Facility of the 16-ft transonic tunnel to determine the performance and turning effectiveness of post-exit yaw vanes installed on two-dimensional convergent-divergent nozzles. One nozzle design that was previously tested was used as a baseline, simulating dry power and afterburning power nozzles at both 0 and 20 degree pitch vectoring conditions. Vanes were installed on these four nozzle configurations to study the effects of vane deflection angle, longitudinal and lateral location, size, and camber. All vanes were hinged at the nozzle sidewall exit, and in addition, some were also hinged at the vane quarter chord (double-hinged). The vane concepts tested generally produced yaw thrust vectoring angles much less than the geometric vane angles, for (up to 8 percent) resultant thrust losses. When the nozzles were pitch vectored, yawing effectiveness decreased as the vanes were moved downstream. Thrust penalties and yawing effectiveness both decreased rapidly as the vanes were moved outboard (laterally). Vane length and height changes increased yawing effectiveness and thrust ratio losses, while using vane camber, and double-hinged vanes increased resultant yaw angles by 50 to 100 percent.

Foley, Robert J.↗
NASA NTRS1989

AFTI/F-111 mission adaptive wing flight test instrumentation overview

The AFTI F-111 Mission Adaptive Wing (MAW) development is described together with the flight test demonstration program. The program developed a smooth variable camber wing and flight control system capable of adjusting the wing's airfoil shape in flight in response to pilot input and flight condition, in order to maximize the aerodynamic efficiency in all areas of flight envelope. This paper examines the instrumentation requirements and the systems implemented and presents the results of flight tests. These results demonstrate the practicality, airworthiness, maintainability, and performance benefits of the smooth variable camber wing system.

Bonnema, Kenneth L.↗
NASA NTRS1991

An assessment of current methods for drag-due-to-lift minimization at supersonic speeds

The estimation and minimization of drag-due-to-lift at supersonic speeds has been examined in this study. Correlations of theory with experimental data are used to assess the applicability and limitations of the linearized theory. The role of leading-edge thrust and the use of twist and camber to develop distributed thrust are also discussed. A semiempirical design and estimation method which takes into account the shortcomings of the linear theory is presented. The use of this method will allow the design of more nearly optimum lifting surfaces and provide an accurate prediction of their level of performance. A preliminary examination is made of the use of an Euler code for estimation of the aerodynamic characteristics of a twisted and cambered wing.

Mann, Michael J.↗
NASA NTRS1991

The turbomachine blading design using S2-S1 approach

The boundary conditions corresponding to the design problem when the blades being simulated by the bound vorticity distribution are presented. The 3D flow is analyzed by the two steps S2 - S1 approach. In the first step, the number of blades is supposed to be infinite, the vortex distribution is transformed into an axisymmetric one, so that the flow field can be analyzed in a meridional plane. The thickness distribution of the blade producing the flow channel striction is taken into account by the modification of metric tensor in the continuity equation. Using the meridional stream function to define the flow field, the mass conservation is satisfied automatically. The governing equation is deduced from the relation between the azimuthal component of the vorticity and the meridional velocity. The value of the azimuthal component of the vorticity is provided by the hub to shroud equilibrium condition. This step leads to the determination of the axisymmetric stream sheets as well as the approximate camber surface of the blade. In the second step, the finite number of blades is taken into account, the inverse problem corresponding to the blade to blade flow confined in each stream sheet is analyzed. The momentum equation implies that the free vortex of the absolute velocity must be tangential to the stream sheet. The governing equation for the blade to blade flow stream function is deduced from this condition. At the beginning, the upper and the lower surfaces of the blades are created from the camber surface obtained from the first step with the assigned thickness distribution. The bound vorticity distribution and the penetrating flux conservation applied on the presumed blade surface constitute the boundary conditions of the inverse problem. The detection of this flux leads to the rectification of the geometry of the blades.

Luu, T. S.↗