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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Computational evaluation of an airfoil with a Gurney flap

A 2D numerical investigation was performed to determine the effect of a Gurney flap on a NACA 4412 airfoil. A Gurney flap is a flat plate on the order of 1 to 3 percent of the airfoil chord length, oriented perpendicular to the airfoil chord line and located at the trailing edge of the airfoil. An incompressible Navier Stokes code, INS2D, was used to calculate the flow field about the airfoil. The fully turbulent results were obtained using the Baldwin-Barth one-equation turbulence model. Gurney flap sizes of 0.5 , 1, 1.25, 1.5, 2, and 3 percent of the airfoil chord were studied. Computational results were compared with experimental results where possible. The numerical solutions show that the Gurney flap increases airfoil lift coefficient with only a slight increase in drag coefficient. Use of a 1.5 percent chord Gurney flap increases the maximum lift coefficient by approximately 0.3 and decreases the angle of attack for a given lift coefficient by more than 3 deg. The numerical solutions exhibit detailed flow structures at the trialing edge and provide a possible explanation for the increased aerodynamic performance.

Jang, Cory S.↗
NASA NTRS1939

Tests in the variable-density wind tunnel of the NACA 23012 airfoil with plain and split flaps

Section characteristics for use in wing design are presented for the NACA 23012 airfoil with plain and split flaps of 20 percent wing chord at a value of the effective Reynolds number of about 8,000,000. The flap deflections covered a range from 60 degrees upward to 75 degrees downward for the plain flap and from neutral to 90 degrees downward for the split flap. The split flap was aerodynamically superior to the plain flap in producing high maximum lift coefficients and in having lower profile-drag coefficients at high lift coefficients.

Abbott, Ira H↗
NASA NTRS1967

The Practical Calculation of the Aerodynamic Characteristics of Slender Finned Vehicles

The basic objective of this thesis is to provide a practical method of computing the aerodynamic characteristics of slender finned vehicles such as sounding rockets, high speed bombs, and guided missiles. The aerodynamic characteristics considered are the normal force coefficient derivative, c(sub N(sub alpha)); center of pressure, bar-X; roll forcing moment coefficient derivative, c(sub l(sub delta)); roll damping moment coefficient derivative, c(sub l(sub p)); pitch damping moment coefficient derivative, c(sub mq); and the drag coefficient, c (sub D). Equations are determined for both subsonic and supersonic flow. No attempts is made to analyze the transonic region. The general configuration to which the relations are applicable is a slender axisymmetric body having three or four fins.

Barrowman, James S.↗
NASA NTRS1983

Experimental wing and canard jet-flap aerodynamics

The effects of upper surface blowing on the aerodynamics of a 1/2-span wing/body/canard configuration are shown. The results expand a data base that is limited at high subsonic Mach numbers (M = 0.6-0.9), data that are needed if computational techniques are to be developed for the complex flowfields generated by jet blowing. At lift coefficients greater than about 1.0, the thrust removed drag coefficient was lower with jet blowing than without jet blowing. This favorable effect increased with increasing jet blowing coefficient, and, for a fixed coefficient, simultaneous wing/canard jet blowing was slightly more effective than blowing either surface alone.

Smeltzer, D. B.↗
NASA NTRS1990

Numerical simulation of an F-16A at angle of attack

The transonic flowfield around an F-16A fighter configuration at a moderate incidence angle is simulated by solving the Navier-Stokes equations on a single-block grid. The numerical solution matches experimental freestream conditions with a mach number of 0.85, 16 degrees angle of attack, and a characteristic Reynolds number of 12.75 million. MacCormack's explicit algorithm is used in conjuction with a local time step and consecutive mesh refinement procedure to accelerate numerical convergence. The Baldwin-Lomax algebraic model provides turbulent closure. Computed surface pressure distributions and the aircraft lift coefficient compare favorably with wind tunnel data. The drag coefficient in the simulation overpredicts the experimental value by 8 percent.

Huband, G. W.↗
NASA NTRS1994

Computation of wind tunnel wall effects for complex models using a low-order panel method

A technique for determining wind tunnel wall effects for complex models using the low-order, three dimensional panel method PMARC (Panel Method Ames Research Center) has been developed. Initial validation of the technique was performed using lift-coefficient data in the linear lift range from tests of a large-scale STOVL fighter model in the National Full-Scale Aerodynamics Complex (NFAC) facility. The data from these tests served as an ideal database for validating the technique because the same model was tested in two wind tunnel test sections with widely different dimensions. The lift-coefficient data obtained for the same model configuration in the two test sections were different, indicating a significant influence of the presence of the tunnel walls and mounting hardware on the lift coefficient in at least one of the two test sections. The wind tunnel wall effects were computed using PMARC and then subtracted from the measured data to yield corrected lift-coefficient versus angle-of-attack curves. The corrected lift-coefficient curves from the two wind tunnel test sections matched very well. Detailed pressure distributions computed by PMARC on the wing lower surface helped identify the source of large strut interference effects in one of the wind tunnel test sections. Extension of the technique to analysis of wind tunnel wall effects on the lift coefficient in the nonlinear lift range and on drag coefficient will require the addition of boundary-layer and separated-flow models to PMARC.

Ashby, Dale L.↗
NASA NTRS1959

Effect of Afterbody-Ejector Configurations on the Performance at Transonic Speeds of a Pylon-Supported Nacelle Model having a Hot-Jet Exhaust

An investigation of several afterbody-ejector configurations on a pylon-supported nacelle model has been completed in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05. The propulsive performance of two nacelle afterbodies with low boattailing and long ejector spacing was compared with a configuration corresponding to a turbojet-engine installation having a highly boattailed afterbody with a short ejector. The jet exhaust was simulated with a hydrogen peroxide turbojet simulator. The angle of attack was maintained at 0 deg, and the average Reynolds number based on body length was 20 x 10(exp 6). The results of the investigation indicated that the configuration with a conical afterbody with smooth transition to a 15 deg boattail angle had large beneficial jet effects on afterbody pressure-drag coefficient and had the best thrust-minus-drag performance of the afterbody-ejector configurations investigated.

Swihart, John M.↗
NASA NTRS1956

Longitudinal and Lateral Stability, Control Characteristics, and Vertical-Tail-Load Measurements for 0.03-Scale Model of the Avro CF-105 Airplane at Mach Number 1.41

An investigation has been made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the aerodynamic characteristics of an 0.03-scale model of the Avro CF-105 airplane. The investigation included the determination of the static longitudinal and lateral stability, the control and the hinge-moment characteristics of the elevator, the aileron, and the rudder, as well as the vertical-tail-load characteristics. The results indicated a minimum drag coefficient of about 0.0270, and a maximum trimmed lift-drag ratio of about 4.25 which occurs at a lift coefficient of 0.16. The directional stability decreased with increasing angle of attack until a region of static instability occurred above an angle of attack of about 9 deg.

Spearman, M. Leroy↗
NASA NTRS1950

Results of the Flight Test of a Dummy of the MX-656 Rocket-Propelled Models

The data obtained from the flight of a simplified (dummy) rocket-propelled model of the MX-656 have been analyzed to determine the booster-model characteristics and the model-alone characteristics up to a Mach number of 1.3. The data indicate that the model-booster combination is satisfactory. The model alone is longitudinally stable i n the Mach number range covered by the test (0.9 to 1.3) with the center of gravity at -15 percent of the mean aerodynamic chord. With the stabilizer setting at 0 deg. the variation of normal-force coefficient with Mach number is not large. The total-drag-coefficient variation with Mach number is not unusual. About 12 percent of the total drag at a Mach number of 1.3 can be attributed to body base drag.

Mitchell, Jesse L.↗
NASA NTRS2015

Skylon Aerodynamics and SABRE Plumes

An independent partial assessment is provided of the technical viability of the Skylon aerospace plane concept, developed by Reaction Engines Limited (REL). The objectives are to verify REL's engineering estimates of airframe aerodynamics during powered flight and to assess the impact of Synergetic Air-Breathing Rocket Engine (SABRE) plumes on the aft fuselage. Pressure lift and drag coefficients derived from simulations conducted with Euler equations for unpowered flight compare very well with those REL computed with engineering methods. The REL coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5, because the engineering estimates did not account for the increasing favorable (in terms of drag and lift coefficients) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown−a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plumeinduced flow separation are other potential risks. The development of an operational reusable launcher from the Skylon concept necessitates the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments.

Mehta, Unmeel↗
NASA NTRS1943

The Effect of Compressibility on the Growth of the Laminar Boundary Layer on Low-Drag Wings and Bodies

The development of the laminar boundary layer in a compressible fluid is considered. Formulas are given for determining the boundary-layer thickness and the ratio of the boundary-layer Reynolds number to the body Reynolds number for airfoils and bodies of revolution. It i s shown that the effect of compressibility will profoundly alter the Reynolds number corresponding to the upper limit of the range of the low-drag coefficients . The available data indicate that for low-drag and high critical compressibility speed airfoils and bodies of revolution, this effect is favorable.

Allen, H. Julian↗
NASA NTRS2018

Steady and Unsteady Excitation of Separated Flow over the NASA Hump Model

Separated flow and its control over the NASA hump model were investigated at subsonic speeds. Three-dimensional, unsteady fluid dynamic simulations were supplemented by wind tunnel measurements. Flow control was implemented by means of spatially distributed discrete jets operating in steady and unsteady modes. The effects of excitation amplitude and frequency were studied numerically and experimentally. Several integral parameters were explored as quality metrics. In addition to the existing ones, two new integral parameters, which are slightly modified versions of the normal force and moment coefficients, were introduced. These two parameters together with pressure drag coefficient were found to be well correlated with the flow control and used in the performance evaluation of different flow control methods including, zero net mass flux actuators, steady suction, sweeping jet actuators, and the currently studied steady and unsteady excitations. For the cases tested, it was found that the unsteady excitation is superior to the steady excitation and slightly better than the sweeping jet actuators whereas the steady suction was found to be the most effective. Although the numerical simulations overpredict the separation bubble, these simulations capture the salient features of flow separation control and hence help us to understand the effect of steady/unsteady excitation on the separated flow.

Koklu, Mehti↗
NASA NTRSDate not supplied

Aerodynamic Design Optimization for Natural Laminar Flow Airfoils

Natural laminar flow technology is a passive laminar flow control (LFC) strategy that seeks to delay the onset of boundary-layer transition (BLT) through shape optimization to reduce the drag of the aerodynamic vehicle. Adjoint-based design optimization for LFC is proposed in an integrated multidisciplinary framework, which includes the computational fluid dynamics (CFD), geometry and grid deformation, and linear stability analysis (LSA) for transition prediction. In particular, the BLT location is predicted using the dual N-factor method that is based on a linear stability theory (LST) eigenvalue problem. The dual N-factor criterion accounts for the amplification of planar Tollmien-Schlichting (TS) and stationary crossflow (CF) boundary-layer instabilities to predict the transition location in three-dimensional boundary-layer flows. The adjoint-based shape optimization procedure is based on an iteratively coupled CFD and LSA methodology to converge the transition location and flow solutions, as well as to calculate the sensitivities of the aerodynamic metrics of interest with respect to the flow and shape design parameters. The RAE 2822 airfoil at 0 and 30 degrees yaw angles, an angle of attack of 0.72 degrees, and subsonic conditions (M∞ = 0.19, Rec = 5.6 × 106 ) are used as baseline configurations for design optimization. The angle of attack and the vertical displacement of free-form-deformation control points are used as design variables to reduce the drag coefficient while reaching a specified lift coefficient. The optimized unswept airfoil designs achieve a 30% drag reduction accompanied by a downstream shift of the transition locations over both suction and pressure sides of the airfoil. The initial design iterations for the swept case also show a favorable trend in the drag reduction with transition delay over both sides.

Transition↗
NASA NTRS2015

Semi-Empirical Prediction of Aircraft Low-Speed Aerodynamic Characteristics

This paper lays out a comprehensive methodology for computing a low-speed, high-lift polar, without requiring additional details about the aircraft design beyond what is typically available at the conceptual design stage. Introducing low-order, physics-based aerodynamic analyses allows the methodology to be more applicable to unconventional aircraft concepts than traditional, fully-empirical methods. The methodology uses empirical relationships for flap lift effectiveness, chord extension, drag-coefficient increment and maximum lift coefficient of various types of flap systems as a function of flap deflection, and combines these increments with the characteristics of the unflapped airfoils. Once the aerodynamic characteristics of the flapped sections are known, a vortex-lattice analysis calculates the three-dimensional lift, drag and moment coefficients of the whole aircraft configuration. This paper details the results of two validation cases: a supercritical airfoil model with several types of flaps; and a 12-foot, full-span aircraft model with slats and double-slotted flaps.

Olson, Erik D.↗
NASA NTRS1977

Application of numerical optimization to the design of supercritical airfoils without drag-creep

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.

Hicks, R. M.↗
NASA NTRS1992

Vortex trapping on a 60 degree delta wing

An experiment was conducted to examine the vortex trapping concept on a 60 deg sweep delta wing at low angles of attack. Using tapered leading edge fences, stable vortices have been successfully trapped. The resulting increase in lift coefficient can be as high as 0.4, which amounts to a 6.5 deg shift in the lift coefficient curve of the delta wing. The corresponding increase in drag coefficient, however, is also significant.

Buchholz, Mark D.↗
NASA NTRS1994

Off-Axis Drag of Dendrite Fragments at Low Reynolds Number

The aim of the present investigation is to characterize the motion of dendrite fragments falling under the influence of gravity in a uniform liquid medium at low Reynolds number. In an earlier study, Zakhem, Weidman and de Groh (1992) reported on the settling speed of model equiaxed dendrite grains released along their axis of symmetry. In this follow-up study uniaxial model dendrite grains were released off-axis to observe and document their motion at different orientations. It was hypothesized that the dendrite models might rotate when released off-axis in which case an attempt would be made to document the ensuing unsteady motion. This latter event turned out to be in fact true: at the small but finite Reynolds numbers that existed, each uniaxial dendrite slowly rotated towards its equilibrium orientation while failing under the influence of gravity. In addition to completing the original goal, we have made use of a beads-on-a shell Stokes flow code to numerically determine the drag coefficient for capsules, i.e.. uniaxial dendrites without arms. The drag on horizontally and vertically falling capsules are reported and compared with measurements.

Weidman, P. D.↗