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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Wall effects in wind tunnels

A synthesis of current trends in the reduction and computation of wall effects is presented. Some of the points discussed include: (1) for the two-dimensional, transonic tests, various control techniques of boundary conditions are used with adaptive walls offering high precision in determining reference conditions and residual corrections. A reduction in the boundary layer effects of the lateral walls is obtained at T2; (2) for the three-dimensional tests, the methods for the reduction of wall effects are still seldom applied due to a lesser need and to their complexity; (3) the supports holding the model of the probes have to be taken into account in the estimation of perturbatory effects.

Chevallier, J. P.↗
NASA NTRS1981

Jet V/STOL wind-tunnel simulation and groundplane effects

Low-speed wind-tunnel testing of V/STOL aircraft concepts to determine the aerodynamic-propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. Powered models are normally used to determine the aerodynamic performance characteristics. This paper examines some of the pretest preparation necessary to define the objectives of an appropriate investigation. Several factors which influence the selection of the model concept and the engine simulator are discussed. In addition, some of the test techniques important for this class of aircraft model are examined. Finally, the paper reviews some of the wind-tunnel wall effects important to this type of aircraft testing with special emphasis on groundplane effects.

Margason, R. J.↗
NASA NTRS1973

Wind tunnel investigation of effects of variations in Reynolds number and leading-edge treatment on the aerodynamic characteristics of an externally blown jet-flap configuration

An investigation has been conducted in a full-scale tunnel to determine the effects of variations in Reynolds number and leading-edge treatment on the aerodynamic characteristics of an externally blown jet-flap transport configuration. The model had a double-slotted trailing-edge flap and was powered by four high-bypass-ratio turbofan engines. Tests were performed by using each of three leading-edge devices (a 30-percent-chord flap and 15- and 25-percent-chord slats) at Reynolds numbers from 0.47 x one million to 1.36 x one million thrust coefficients up to 3.5. The use of a 25-percent-chord slat was found to be more effective than a 15-percent-chord slat or a 30-percent-chord flap in extending the stall angle of attack and in minimizing the loss of lift after the stall. The large slat was also effective in reducing the rolling moments that occurred when the engine-out wing stalled first.

Parlett, L. P.↗
NASA NTRS1924

The effect of wind tunnel turbulence upon the forces measured on models

1. Reasons for inquiry: The tests were undertaken to find the effect of turbulence in the air stream upon the lift and drag forces measured on models in the four-foot wind tunnel at the Massachusetts Institute of Technology. 2. Range of investigation: Maximum lifts and minimum drags were measured on Gottingen-387 and R.A.F.-15 airfoils, minimum drag on a streamlined strut, and the static pressure gradients for different conditions of turbulence. 3. Results and further developments: The results show that the scale of the turbulence (as defined in this report) has a marked effect upon the measured forces on models tested in the tunnel as well as on the pressure gradient, and it is recommended that further investigation of the phenomena be made with the aid of smoke and small wind vanes.

Lepage, W L↗
NASA NTRS1948

Tests of a Triangular Wing of Aspect Ratio 2 in the Ames 12-foot Pressure Wind Tunnel III : the Effectiveness and Hinge Moments of a Skewed Wing-tip Flap

Results of wind-tunnel tests of a semispan model of a triangular wing of aspect ratio 2 with a skewed wing-tip flap are presented. Lift, drag, pitching-moment, and hinge-moment data are included for subsonic Mach numbers up to 0.95. The flap showed extremely high hinge moments and low effectiveness as a longitudinal control. Although less affected by compressibility, this flap is indicated to be inferior to a constant-chord flap when applied to this triangular wing.

MACH NUMBER EFFECTS - COMPLETE WINGS↗
NASA NTRS1972

Numerical calculation of boundary-induced interference in slotted or perforated wind tunnels including viscous effects in slots

A numerical method is presented for calculating the incompressible boundary-induced interference in wind tunnels of rectangular cross section with slotted or perforated walls. The method includes a wall representation which is capable of satisfying a generalized homogeneous boundary condition including the effects of viscosity within the slots. The effects of viscosity in the slots are found to be very significant. The method allows for a variation in the boundary conditions along the tunnel walls. The model can be any configuration and can be located anywhere in the test section. The interference can be computed at any point in the test section.

Keller, J. D.↗
NASA NTRS1947

Wind-Tunnel Investigation of Effects of Unsymmetrical Horizontal-Tail Arrangements on Power-on Static Longitudinal Stability of a Single-Engine Airplane Model

A wind-tunnel investigation has been made to determine the effects of unsymmetrical horizontal-tail arrangements on the power-on static longitudinal stability of a single-engine single-rotation airplane model. Although the tests and analyses showed that extreme asymmetry in the horizontal tail indicated a reduction in power effects on longitudinal stability for single-engine single-rotation airplanes, the particular "practical" arrangement tested did not show marked improvement. Differences in average downwash between the normal tail arrangement and various other tail arrangements estimated from computed values of propeller-slipstream rotation agreed with values estimated from pitching-moment test data for the flaps-up condition (low thrust and torque) and disagreed for the flaps-down condition (high thrust and torque). This disagreement indicated the necessity for continued research to determine the characteristics of the slip-stream behind various propeller-fuselage-wing combinations. Out-of-trim lateral forces and moments of the unsymmetrical tail arrangements that were best from consideration of longitudinal stability were no greater than those of the normal tail arrangement.

AIRFOILS - NACA 65(216-012(MOD ) - CHAR & MOMENTS ↗
NASA NTRS1985

Low-speed stability and control wind-tunnel investigations of effects of spanwise blowing on fighter flight characteristics at high angles of attack

The effects of spanwise blowing on two configurations representative of current fighter airplanes were investigated. The two configurations differed only in wing planform, with one incorporating a trapezoidal wing and the other a 60 delta wing. Emphasis was on determining the lateral-directional characteristics, particularly in the stall/departure angle-of-attack range; however, the effects of spanwise blowing on the longitudinal aerodynamics were also determined. The-tunnel tests included measurement of static force and forced-oscillation aerodynamic data, visualization of the airflow changes created by the spanwise blowing, and free-flight model tests. The effects of blowing rate, chordwise location of the blowing ports, asymmetric blowing, and blowing on the conventional aerodynamic control characteristics were investigated. In the angle-of-attack regions in which the spanwise blowing substantially improved the wing upper-surface flow field (i.e., provided reattachment of the flow aft of the leading-edge vortex), improvements in both static and dynamic lateral-directional stability were observed. Blowing effects on stability could be proverse or adverse depending on blowing rate, blowing port loaction, and wing planform. Free-flight model tests of the trapezoidal wing confirmed the beneficial effects of spanwise blowing measured in the static and dynamic force tests.

Satran, D. R.↗
NASA NTRS1996

Hypersonic Boundary Layer Stability Experiments in a Quiet Wind Tunnel with Bluntness Effects

Hypersonic boundary layer measurements over a flared cone were conducted in a Mach 6 quiet wind tunnel at a freestream unit Reynolds number of 2.82 million/ft. This Reynolds number provided laminar-to-transitional flow over the cone model in a low-disturbance environment. Four interchangeable nose-tips, including a sharp-tip, were tested. Point measurements with a single hot-wire using a novel constant voltage anemometer were used to measure the boundary layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the transitional state of the boundary layer and to identify instability modes. Results suggest that second mode disturbances were the most unstable and scaled with the boundary layer thickness. The second mode integrated growth rates compared well with linear stability theory in the linear stability regime. The second mode is responsible for transition onset despite the existence of a second mode subharmonic. The subharmonic disturbance wavelength also scales with the boundary layer thickness. Furthermore, the existence of higher harmonics of the fundamental suggests that nonlinear disturbances are not associated with 'high' free stream disturbance levels. Nose-tip radii greater than 2.7% of the base radius completely stabilized the second mode.

Lachowicz, Jason T.↗
NASA NTRS1938

Correction of Profile-Drag Results from Variable-Density Tunnel and the Effect on the Choice of Wing-Section Thickness

Profile-drag coefficients published from tests in the N.A.C.A. variable-density tunnel (Technical Reports Nos. 460, 537, 586, and 610, references 1 to 4) have tended to appear high as compared with results from the N.A.C.A. full-scale tunnel (Technical Report No. 530, reference 5) and from foreign sources (references 6 to 8). Such discrepancies were considered in Technical Report No. 586, and corrections for turbulence and tip effects were derived that tended to reduce the profile-drag coefficients, particularly for the thicker airfoils. The corrected profile-drag coefficients, designated by the lower-case symbol cdo as contrasted with the older CDO, have been employed in the airfoil reports published since Technical Report No. 460, but even these corrected results continued to appear high, particularly for the thicker sections. The important practical result is that a smaller increase of drag with airfoil thickness is indicated, which may be of primary importance to the airplane designer in choosing the optimum airfoil sections for actual wings. Further investigations of this subject were, of course, undertaken, one of the most important being an investigation of three symmetrical sections N.A.C A. 0009, 0012, and 0018 under conditions of low turbulence in the full-scale tunnel. Preliminary results from this investigation also indicate a smaller increase in drag with airfoil thickness than the results from the variable-density tunnel. Furthermore, comparative tests made in the two tunnels by applying strings to the surface of the N.A.C.A. 0012 airfoil to move the transition point to a predetermined position indicated that the effective reynolds Number concept would account approximately for the drag as affected by the position of transition from laminar to turbulent flow in the boundary layer.

Jacobs, Eastman N.↗
NASA NTRS1974

Continued investigations in the NAAL low speed wind tunnel into the effects of the air breathing propulsion system on orbiter subsonic stability and control characteristics (OA62A)

Experimental aerodynamic investigations were conducted on a stingmounted 0.0405-scale representation (model 43-0) of the 140A/B Space Shuttle Orbiter in a Low Speed Wind Tunnel. The NASA designation for this test was 0A62A. The primary test objective was to continue studies, initiated on tests 0A16 and 0A71A and 0A71C, in optimizing the air breathing propulsion system (ABPS) and investigating the aerodynamic effects of various nacelle number/location configurations on the orbiter stability and control characteristics. Orbiter stability and control characteristics, both with and without ABPS, were investigated at elevon deflections of 0, + or -5, + or -19, + or -5, and -20 deg; aileron deflections of 0 and 10 deg (about 0 deg elevon); and rudder deflections of 0, -7.5, and -15 deg. Aerodynamic force and moment data was measured in the body axis system by a 2.5-inch task type internal balance. The model was sting supported through the base region with a nominal angle of attack range of -4 to 30 deg. Yaw polars were recorded over the beta range of -10 to 10 deg at fixed angles of attack of 0, 5, 10, and 15 deg.

Mennell, R.↗
NASA NTRS1939

Wind-tunnel Investigation of Effect of Yaw on Lateral-stability Characteristics I : Four N.A.C.A. 23012 Wings of Various Plan Forms with and Without Dihedral

Four N.A.C.A. 23012 wings were tested at several angles of yaw in the N.A.C.A. 7- by 10-foot wind tunnel. All the wings have rounded tips and, in plan form, one is rectangular and the others are tapered 3:1 with various amounts of sweep. Each wing was tested with two amounts of dihedral and with partial-span split flaps. The coefficients of lift, drag, and pitching moment are given for all the models at zero yaw. The coefficients of rolling moment, yawing moment, and side force are given for the rectangular wing at all values of yaw tested. The rate of change in the coefficients with angle of yaw is given in convenient form for stability calculations.

Bamber, M J↗
NASA NTRS1976

Effect of slot width on transition and noise attenuation of a sound shield panel for supersonic wind tunnels

The effect of slot width on transition and noise attenuation of a planar rod wall sound shield has been investigated at Mach 6. Boundary layer laminarization was obtained by induced suction through adjustable slots or gaps between the rods. Transition Reynolds numbers increased significantly with increasing gap width. The local mean flow field above the suction panel rod array was uniform. Important factors for design and performance of a supersonic wind-tunnel sound shield are rod diameter, gap-to-rod diameter ratio, surface finish of the rods, and leading-edge configuration.

Harvey, W. D.↗
NASA NTRS1982

Wind-tunnel investigation of effects of wing-leading-edge modifications on the high angle-of-attack characteristics of a T-tail low-wing general-aviation aircraft

Exploratory tests have been conducted in the NASA-Langley Research Center's 12-Foot Low-Speed wind Tunnel to evaluate the application of wing-leading-edge devices on the stall-departure and spin resistance characteristics of a 1/6-scale model of a T-tail general-aviation aircraft. The model was force tested with an internal strain-gauge balance to obtain aerodynamic data on the complete configuration and with a separate wing balance to obtain aerodynamic data on the outer portion of the wing. The addition of the outboard leading-edge droop eliminated the abrupt stall of the windtip and maintained or increased the resultant-force coefficient up to about alpha = 32 degrees. This change in slope of the resultant-force coefficient curve with angle of attack has been shown to be important for eliminating autorotation and for providing spin resistance.

White, E. R.↗
NASA NTRS1959

Wind-Tunnel Measurements of Effect of Dive-Recovery Flaps at Transonic Speeds on Models of a Seaplane and a Transport

The effects of wing-lower-surface dive-recovery flaps on the aero- dynamic characteristics of a transonic seaplane model and a transonic transport model having 40 deg swept wings have been investigated in the Langley 16-foot transonic tunnel. The seaplane model had a wing with an aspect ratio of 5.26, a taper ratio of 0.333, and NACA 63A series airfoil sections streamwise. The transport model had a wing with an aspect ratio of 8, a taper ratio of 0.3, and NACA 65A series airfoil sections perpendicular to the quarter-chord line. The effects of flap deflection, flap longitudinal location, and flap sweep were generally investigated for both horizontal-tail-on and horizontal-tail-off configurations. Model force and moment measurements were made for model angles of attack from -5 deg to 14 deg in the Mach number range from 0.70 to 1.075 at Reynolds numbers of 2.95 x 10(exp 6) to 4.35 x 10(exp 6). With proper longitudinal location, wing-lower-surface dive-recovery flaps produced lift and pitching-moment increments that increased with flap deflection. For the transport model a flap located aft on the wing proved to be more effective than one located more forward., both flaps having the same span and approximately the same deflection. For the seaplane model a high horizontal tail provided added effectiveness for the deflected-flap configuration.

Heath, Atwood R., Jr.↗