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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 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Combined, nonlinear aerodynamic and structural method for the aeroelastic design of a three-dimensional wing in supersonic flow

An iterative procedure for the static aeroelastic design of a flexible wing at supersonic speeds has been developed. The procedure combines a nonlinear, full-potential solver (NCOREL) with an equivalent plate structural analysis method. The NCOREL method yields significantly improved aerodynamic estimates compared to linear theory. The equivalent plate structural analysis method demonstrates an order of magnitude reduction in computer memory and execution time compared to finite-element methods. A highly swept wing is analyzed at high lift using this aeroelastic procedure. The results indicate that the wing deforms favorably due to aerodynamic loading and, consequently, that the inviscid drag levels do not vary at the required lift coefficient although the angle of attack varies significantly. A sensitivity analysis of the type required for optimization studies was also performed with the aeroelastic design procedure.

Pittman, J. L.↗
NASA NTRS1985

Angle-of-attack estimation for analysis of CAT encounters

Recent studies of clear-air turbulence (CAT) encounters involving wide-body airliners have been based upon flight-path wind estimates made by analyzing digital flight-data-recorder (DFDR) records and radar records. Such estimates require a time history of the aircraft angle of attack, a record that is not usually included in the DFDR measurement set. This paper describes a method for reconstructing angle of attack that utilizes available flight record and aircraft-specific information associated with an aerodynamic model of the lift coefficient. Results from two wide-body incidents in which vane measurements of angle of attack were recorded show good agreement between measured and calculated time histories. This research has been performed in cooperation with the National Transportation Safety Board to provide a better understanding of the CAT phenomenon.

Bach, R. E., Jr.↗
NASA NTRS1984

Supersonic, nonlinear, attached-flow wing design for high lift with experimental validation

Results of the experimental validation are presented for the three dimensional cambered wing which was designed to achieve attached supercritical cross flow for lifting conditions typical of supersonic maneuver. The design point was a lift coefficient of 0.4 at Mach 1.62 and 12 deg angle of attack. Results from the nonlinear full potential method are presented to show the validity of the design process along with results from linear theory codes. Longitudinal force and moment data and static pressure data were obtained in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.58, 1.62, 1.66, 1.70, and 2.00 over an angle of attack range of 0 to 14 deg at a Reynolds number of 2.0 x 10 to the 6th power per foot. Oil flow photographs of the upper surface were obtained at M = 1.62 for alpha approx. = 8, 10, 12, and 14 deg.

Pittman, J. L.↗
NASA NTRS1984

Numerical methods and a computer program for subsonic and supersonic aerodynamic design and analysis of wings with attainable thrust considerations

This paper describes methodology and an associated computer program for the design of wing lifting surfaces with attainable thrust taken into consideration. The approach is based on the determination of an optimum combination of a series of candidate surfaces rather than the more commonly used candidate loadings. Special leading-edge surfaces are selected to provide distributed leading-edge thrust forces which compensate for any failure to achieve the full theoretical leading-edge thrust, and a second series of general candidate surfaces is selected to minimize drag subject to constraints on the lift coefficient and, if desired, on the pitching moment coefficient. A primary purpose of the design approach is the introduction of attainable leading-edge thrust considerations so that relatively mild camber surfaces may be employed in the achievement of aerodynamic efficiencies comparable to those attainable if full theoretical leading-edge thrust could be achieved. The program provides an analysis as well as a design capability and is applicable to both subsonic and supersonic flow.

Carlson, H. W.↗
NASA NTRS1986

Forward-swept wing configuration designed for high maneuverability by use of a transonic computational method

A transonic computational analysis method and a transonic design procedure have been used to design the wing and the canard of a forward-swept-wing fighter configuration for good transonic maneuver performance. A model of this configuration was tested in the Langley 16-Foot Transonic Tunnel. Oil-flow photographs were obtained to examine the wind flow patterns at Mach numbers from 0.60 to 0.90. The transonic theory gave a reasonably good estimate of the wing pressure distributions at transonic maneuver conditions. Comparison of the forward-swept-wing configuration with an equivalent aft-swept-wing-configuration showed that, at a Mach number of 0.90 and a lift coefficient of 0.9, the two configurations have the same trimmed drag. The forward-swept wing configuration was also found to have trimmed drag levels at transonic maneuver conditions which are comparable to those of the HiMAT (highly maneuverable aircraft technology) configuration and the X-29 forward-swept-wing research configuration. The configuration of this study was also tested with a forebody strake.

Mann, M. J.↗
NASA NTRS1987

Subsonic maneuver capability of a supersonic cruise fighter wing concept

A theoretical and experimental investigation was conducted of the subsonic maneuver capability of a fighter wing concept designed for supersonic cruise. To improve the subsonic maneuver capability, the wing utilized full-span leading- and trailing-edge flaps that were designed with the aid of a subsonic-analysis computer program. Wind-tunnel tests were made at Mach numbers of 0.3, 0.5, and 0.7. Force and moment data obtained were compared with theoretical predictions of Mach 0.5 from two subsonic-analysis computer programs. The two theoretical programs gave a good prediction of the lift and drag characteristics but only a fair prediction of the pitching moment. The experimental results of this study show that with the proper combination of leading- and trailing-edge flap deflections, a suction parameter of nearly 90 percent can be attained at a Mach number of 0.5 and a lift coefficient of 0.73; this is a three-fold improvement from 30 percent for the basic wing.

Riebe, Gregory D.↗
NASA NTRS1985

Analysis of the free-tip rotor wind-tunnel test results

The results from a wind tunnel test of a small scale free-tip rotor are analyzed. The free-tip rotor has blade tips that are free to weathervane into the tip's relative wind, thus producing a more uniform lift around the azimuth. The free-tip assembly, which includes the controller, functioned flawlessly throughout the test. In a test of the free-tip's response after passing through a vertical air jet, the tip pitched freely and in a controlled manner. Analysis of the tip's response characteristics showed the free-tip system's damped natural frequency to be 5.2 per rev. Tip pitch angle responses to the local airstream are presented for an advance-ratio range of 0.1 to 0.397 and for a solidity weighted rotor lift coefficient range of 0.038 to 0.092. Harmonic analysis of the responses showed a dominance by the first harmonic. As a result of the tip being free, forward flight power requirements were reduced by 8% or more. More power reduction was recorded for high thrust conditions. In addition to the power reduction, flatwise blade bending moments were reduced by as much as 30% at the inboard blade stations.

Stroub, Robert H.↗
NASA NTRS1986

Effect of an upstream wake on a pusher propeller

An analytical/computational study has been conducted to predict the effect of an upstream wake on an operating propeller. The upstream wing/pylon was modeled by a constant chord wing of NACA 0012 sections and was placed at a variable distance (0.1 - 0.3 chord) upstream of a scaled model propeller (SR-2). The wake model was a similarity formulation. The periodic behavior ofthe flow during the passage through the wake was formulated in terms of time-dependent variation of each blade section's angle of attack. It was found that the final expressions for the unsteady pressure distribution on each blade section are periodic and that the unsteady circulation and lift coefficieents exhibit a hysteresis loop.

Takallu, M. A.↗
NASA NTRS1987

Wind-tunnel test results of airfoil modifications for the EA-6B

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.

Sewall, W. G.↗
NASA NTRS1987

EA-6B high-lift wing modifications

NASA-Langley has accomplished the computational design and experimental verification of EA-6B aircraft wing modifications for improved high lift capability. The modifications are comparatively simple, and attempt to improve low speed high lift performance while maintaining high speed cruise efficiency. Several two- and three-dimensional low speed and transonic computational techniques were employed, together with extensive wind tunnel tests. The modified inboard and outboard edge slat/flap system sections yielded efficiency improvements that were verified by three-dimensional wind tunnel experiments to amount to an 11-percent wing-body lift coefficient enhancement at low speed.

Waggoner, E. G.↗
NASA NTRS1984

An Experimental Evaluation of Advanced Rotorcraft Airfoils in the NASA Ames Eleven-foot Transonic Wind Tunnel

Five full scale rotorcraft airfoils were tested in the NASA Ames Eleven-Foot Transonic Wind Tunnel for full scale Reynolds numbers at Mach numbers from 0.3 to 1.07. The models, which spanned the tunnel from floor to ceiling, included two modern baseline airfoils, the SC1095 and SC1094 R8, which have been previously tested in other facilities. Three advanced transonic airfoils, designated the SSC-A09, SSC-A07, and SSC-B08, were tested to confirm predicted performance and provide confirmation of advanced airfoil design methods. The test showed that the eleven-foot tunnel is suited to two-dimensional airfoil testing. Maximum lift coefficients, drag coefficients, pitching moments, and pressure coefficient distributions are presented. The airfoil analysis codes agreed well with the data, with the Grumman GRUMFOIL code giving the best overall performance correlation.

Flemming, Robert J.↗
NASA NTRS1985

Wind tunnel results for a high-speed, natural laminar-flow airfoil designed for general aviation aircraft

Two dimensional wind tunnel tests were conducted on a high speed natural laminar flow airfoil in both the Langley 6 x 28 inch Transonic Tunnel and the Langley Low Turbulence Pressure Tunnel. The test conditions consisted of Mach numbers ranging from 0.10 to 0.77 and Reynolds numbers ranging from 3 x 1 million to 11 x 1 million. The airfoil was designed for a lift coefficient of 0.20 at a Mach number of 0.70 and Reynolds number of 11 x 1 million. At these conditions, laminar flow would extend back to 50 percent chord of the upper surface and 70 percent chord of the lower surface. Low speed results were also obtained with a 0.20 chord trailing edge split flap deflected 60 deg.

Sewall, William G.↗
NASA NTRS1988

Comparison of pressure distributions on model and full-scale NACA 64-621 airfoils with ailerons for wind turbine application

The aerodynamic similarity between a small (4-inch chord) wind tunnel model and a full-scale wind turbine blade (24-foot tip section with a 36-inch chord) was evaluated by comparing selected pressure distributions around the geometrically similar cross sections. The airfoils were NACA 64-621 sections, including trailing-edge ailerons with a width equal to 38 percent of the airfoil chord. The model airfoil was tested in the OSU 6- by 12-inch High Reynolds Number Wind Tunnel; the full-scale blade section was tested in the NASA Langley Research Center 30- by 60-foot Subsonic Wind Tunnel. The model airfoil contained 61 pressure taps connected by embedded tubes to pressure transducers. A belt containing 29 pressure taps was fixed to the full-scale section at midspan to obtain surface pressure data. Lift coefficients were obtained by integrating pressures, and corrections were made for the 3-D effects of blade twist and downwash in the blade tip section. The results of the two different experimental methods correlated well for angles of attack from minus 4 to 36 degrees and aileron reflections from 0 to 90 degrees.

Gregorek, G. M.↗
NASA NTRS1985

Parametric study of supersonic STOVL flight characteristics

A number of different control devices and techniques are evaluated to determine their suitability for increasing the short takeoff performance of a supersonic short-takeoff/vertical landing (STOVL) aircraft. Analysis was based on a rigid-body mathematical model of the General Dynamics E-7, a single engine configuration that utilizes ejectors and thrust deflection for propulsive lift. Alternatives investigated include increased static pitch, the addition of a close-coupled canard, use of boundary layer control to increase the takeoff lift coefficient, and the addition of a vectorable aft fan air nozzle. Other performance studies included the impact of individual E-7 features, the sensitivity to ejector performance, the effect of removing the afterburners, and a determination of optional takeoff and landing transition methods. The results pertain to both the E-7 and other configurations. Several alternatives were not as well suited to the E-7 characteristics as they would be to an alternative configuration, and vice versa. A large amount of supporting data for each analysis is included.

Rapp, David C.↗
NASA NTRS1988

Euler equation computations for the flow over a hovering helicopter rotor

A numerical solution technique is developed for computing the flow field around an isolated helicopter rotor in hover. The flow is governed by the compressible Euler equations which are integrated using a finite volume approach. The Euler equations are coupled to a free wake model of the rotary wing vortical wake. This wake model is incorporated into the finite volume solver using a prescribed flow, or perturbation, technique which eliminates the numerical diffusion of vorticity due to the artificial viscosity of the scheme. The work is divided into three major parts: (1) comparisons of Euler solutions to experimental data for the flow around isolated wings show good agreement with the surface pressures, but poor agreement with the vortical wake structure; (2) the perturbation method is developed and used to compute the interaction of a streamwise vortex with a semispan wing. The rapid diffusion of the vortex when only the basic Euler solver is used is illustrated, and excellent agreement with experimental section lift coefficients is demonstrated when using the perturbation approach; and (3) the free wake solution technique is described and the coupling of the wake to the Euler solver for an isolated rotor is presented. Comparisons with experimental blade load data for several cases show good agreement, with discrepancies largely attributable to the neglect of viscous effects. The computed wake geometries agree less well with experiment, the primary difference being that too rapid a wake contraction is predicted for all the cases.

Roberts, Thomas Wesley↗
NASA NTRS1988

High-aspect-ratio wings

High-aspect-ratio aircraft include most transport aircraft such as commercial and military transports, business aircraft, and cargo aircraft. Generally, these types of aircraft are designed to cruise over a narrow range of lift coefficients and Mach numbers in the performance of their mission. Emphasis is therefore placed on the cruise performance of transport aircraft and every effort is made to obtain accurate wind-tunnel data to use as a basis for prediction of full-scale cruise performance. However, off-cruise performance is also important and methods were developed for extrapolating wind-tunnel data on buffet and flutter at transonic speed. Transport-type aircraft were tested extensively in various wind tunnels around the world and many different test techniques were developed to simulate higher Reynolds numbers. Methods developed for one tunnel may not be applicable to another tunnel because of differences in size, Reynolds number capability, running time, and test objectives. Many of the methods of boundary-layer control developed in two-dimensional airfoil testing can be applied in tests of transport configurations, but sometimes the three-dimensional flow fields that develop on tranpsort aircraft can make application of the two-dimensional methods difficult or impossible. The discussion is intended to be a representative, but not exhaustive, survey of the various methods of high Reynolds number simulation in the testing of high-aspect-ratio aircraft.

Peterson, John B., Jr.↗
NASA NTRS1986

Test results at transonic speeds on a contoured over-the-wing propfan model

A semispan wing/body model with a powered highly loaded propeller has been tested to provide data on the propulsion installation drag of advanced propfan-powered aircraft. The model had a supercritical wing with a contoured over-the-wing nacelle. It was tested in the Ames Research Center's (ARC) 14-foot Transonic Wind Tunnel at a total pressure of 1 atm. The test was conducted at angles of attack from -0.5 to 4 deg at Mach numbers ranging from 0.6 to 0.8. The test objectives were to determine propeller performance, exhaust jet effects, propeller slipstream interference drag, and total powerplant installation drag. Test results indicated a total powerplant installation drag of 82 counts (0.0082) at a Mach number of 0.8 and a lift coefficient of 0.5, which is approximately 29 percent of a typical airplane cruise drag.

Levin, Alan D.↗
NASA NTRS1987

Optimal trajectories for aeroassisted, noncoplanar orbital transfer. II - LEO-to-LEO transfer

Both classical and minimax problems of optimal control arising in the study of noncoplanar, aeroassisted orbital transfer are considered and are illustrated with the example of LEO-to-LEO transfer. Trajectory control is achieved by modulation of the lift coefficient and the angle of bank. Problems considered include the minimization of the energy required for orbital transfer, maximization of the flight time during the atmospheric portion of the trajectory, and minimization of the peak heating rate. The near-grazing solution is found to be a good compromise between energy and heating requirements.

Miele, A.↗