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An investigation has been made to determine the effect of control-surface profile modifications on the aerodynamic characteristics of an NACA low-drag airfoil equipped with a 0.20-chord and a 0.15-chord aileron. Tab characteristics have been obtained for 0.20-aileron chord tabs on two of the 0.20-chord ailerons. Basic data are presented from which the effect of tabs can be calculated for specific cases. The data are sufficient for the solution of problems of fixed tabs with a differential linkage, as well as simple and spring-linked balancing tabs.
Data on the aerodynamic excitation of first-order vibration occurring in a representative three-blade propeller having its thrust axis inclined to the airstream at angles of 0 degree, 4.55 degrees, and 9.80 degrees are included in this report. For several representative conditions the aerodynamic excitation has been computed and compared with the measured values. Blade stresses also were measured to permit the evaluation of the blade stress resulting from a given blade aerodynamic excitation. It was concluded that the section aerodynamic exciting force of a pitched propeller may be computed accurately at low rotational speeds. As section velocities approach the speed of sound, the accuracy of computation of section aerodynamic exciting force is not always so satisfactory.
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Data on the aerodynamic excitation of first-order vibration occurring in a representative three-blade propeller having its thrust axis inclined to the air stream at angles of 0 degrees, 4.55 degrees, and 9.8 degrees are included in this paper. For several representative conditions the aerodynamic excitation has been computed and compared with the measured values. Blade stresses also were measured to permit the evaluation of the blade stress resulting from a given blade aerodynamic excitation. It was concluded that the section aerodynamic exciting force of a pitched propeller may be computed accurately at low rotational speeds. As section velocities approach the speed of sound, the accuracy of computation of section aerodynamic exciting force is not always so satisfactory. First-order blade vibratory stresses were computed with satisfactory accuracy from untilted-propeller loading data. A stress prediction which assumes a linear relation between first-order vibratory stress and the product of pitch angle and dynamic pressure and which is based on stresses at low rotational speeds will be conservative when the outer portions of the blade are in the transonic and low supersonic speed range.
A research model of an airplane with a configuration suitable for supersonic flight was tested at transonic speeds in order to establish the effects on longitudinal and lateral stability of certain changes in both wing sweep and height of the horizontal tail. Two wings of aspect ratio 3 and taper ratio 0.15, one having the quarter-chord line swept back 30 deg and the other 45 deg, were each tested with the horizontal tail of the model in a low and in a high position. One configuration was also tested with fuselage strakes. The tests were made at Mach numbers from 0.60 to 1.17 and Reynolds numbers from 1.9 x 10(exp 6) to 2.6 x 10(exp 6). The results indicated that a low horizontal-tail position (below the wing-chord plane) gave positive longitudinal stability for the model for all angles of attack used (angles of attack up to 24 deg); whereas, a higher tail position (above the wing-chord plane) resulted in a large reduction in stability at moderate angles of attack. With the higher horizontal tail, the 30 deg-swept-wing model had somewhat more stability than the 45 deg-swept-wing model at subsonic Mach numbers. With the lower tail, the 45 deg-swept-wing model had slightly more stability at all Mach numbers. The model with the 30 deg swept wing had greater directional stability with the tail in the higher rather than the lower position, but the opposite was true for the 45 deg-swept-wing model. The directional stability decreased sharply at high angles of attack; this characteristic was alleviated by the use of fuselage strakes which, however, proved to be detrimental to the longitudinal stability of the model tested.
The experimental wing buffet response of a transport-type airplane model with and without wing bodies, fences, flaps, and a fuselage addition has been investigated at Mach numbers from 0.20 to 1.03. The wing had NACA 64A-series airfoil sections inclined 5 degrees to the free-stream direction. The quarter-chord line of the wing was swept back 45 degrees, the aspect ratio was 7, the taper ratio was 0.3, and the thickness ratio varied from 0.115 at the root to 0.074 at the midsemispan and was constant from that station to the tip. The wing was twisted and cambered for a design lift coefficient of 0.3. The results of the investigation indicated that a marked reduction of buffet intensity and a delay of buffet onset at transonic speeds were achieved by the addition to the wing of special bodies designed to reduce shock-induced separation. The further addition of wing fences and wing trailing-edge flaps deflected 30 degrees increased the lift coefficients at which low-speed stall buffeting occurred. An addition to the fuselage near the upper forward portion produced no consistent change in the buffet characteristics.
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This slide presentation reviews a method for the assessment of three-dimensional conduction effects during test in a Aerothermal Wind Tunnel. The test objectives were to duplicate and extend tests that were performed during the 1960's on thermal conduction on proturberance on a flat plate. Slides review the 1D versus 3D conduction data reduction error, the analysis process, CFD-based analysis, loose coupling method that simulates a wind tunnel test run, verification of the CFD solution, Grid convergence, Mach number trend, size trends, and a Sumary of the CFD conduction analysis. Other slides show comparisons to pretest CFD at Mach 1.5 and 2.16 and the geometries of the models and grids.
A low speed wind tunnel test was conducted to assess the effects of the larger JT8D refan nacelles on the stability and control characteristics of the DC-9-30, with emphasis on the deep stall regime. Deep stall pitching moment and elevator hinge moment data, and low angle of attack tail-on and tail-off pitching moment data are presented. The refan nacelle was tested in conjunction with various pylons of reduced span relative to the production DC-9-30 pylon. Also, a horizontal tail that was larger than the production tail was tested. The data show that the refan installation has a small detrimental effect on the DC-9-30 deep stall recovery capability, that recovery characteristics are essentially independent of pylon span, and that the larger horizontal tail significantly increases recovery margins. The deep stall characteristics with the refan installation, within the range of pylon spans tested, are acceptable with no additional design changes anticipated.
Wind tunnel tests were made for spheres of various sizes over a range of Mach numbers and Reynolds numbers. The results indicated some conditions where the drag was affected by changes in the afterbody pressure due to a shock reflection from the tunnel wall. This effect disappeared when the Mach number was increased for a given sphere size or when the sphere size was decreased for a given Mach number. Drag measurements and Schlieren photographs show the possibility of obtaining inaccurate data when tests are made with a sphere too large for the test section size and Mach number. Tests were also made of an oblate spheroid. The results indicated a region at high Mach numbers where inherent positive static stability might occur with the oblate-face forward. The drag results are compared with those for a sphere and those for various other shapes. The drag results for the oblate spheriod and the sphere are also compared with some calculated results.
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Analytical investigation of reentry behavior of flying wind tunnel test vehicle with some effects of threshold and torque level of roll-rate control system
Rayleigh scattering, a nonintrusive measurement technique for the measurement of density in a hypersonic wind tunnel, is under investigation at Wright Laboratory's Mach 6 wind tunnel. Several adverse effects, i.e., extraneous scatter off walls and windows, hinder Rayleigh scattering measurements. Condensation and clustering of flow constituents also present formidable obstacles. Overcoming some of these difficulties, measurements have been achieved while the Mach 6 test section was pumped down to a vacuum, as well as for actual tunnel operation for various stagnation pressures at fixed stagnation temperatures. Stagnation pressures ranged from 0.69 MPa to 6.9 MPa at fixed stagnation temperatures of 511, 556, and 611 K. Rayleigh scatter results show signal levels much higher than expected for molecular scattering in the wind tunnel. Even with higher than expected signals, scattering measurements have been made in the flowfield of an 8-degree half-angle blunt nose cone with a nose radius of 1.5 cm.
An investigation has been conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel in order to further the development of semi-span testing capabilities. A twin engine, energy efficient transport (EET) model with a four-element wing in a takeoff configuration was used for this investigation. Initially a full span configuration was tested and force and moment data, wing and fuselage surface pressure data, and fuselage boundary layer measurements were obtained as a baseline data set. The semi-span configurations were then mounted on the wind tunnel floor, and the effects of fuselage standoff height and shape as well as the effects of the tunnel floor boundary layer height were investigated. The effectiveness of tangential blowing at the standoff/floor juncture as an active boundary-layer control technique was also studied. Results indicate that the semi-span configuration was more sensitive to variations in standoff height than to variations in floor boundary layer height. A standoff height equivalent to 30 percent of the fuselage radius resulted in better correlation with full span data than no standoff or the larger standoff configurations investigated. Undercut standoff leading edges or the use of tangential blowing in the standoff/ floor juncture improved correlation of semi-span data with full span data in the region of maximum lift coefficient.
Cone-cylinder models tested in supersonic wind tunnel to evaluate effect of wall reflected disturbances on model pressure distributions
Report presents the results of tests of a 1/10-scale model of the XN2Y-1 airplane tested in the NACA 5-foot vertical wind tunnel in which the six components of forces and moments were measured. The model was tested in 17 attitudes in which the full-scale airplane had been observed to spin, in order to determine the effects of scale, tunnel, and interference. In addition, a series of tests was made to cover the range of angles of attack, angles of sideslip, rates of rotation, and control setting likely to be encountered by a spinning airplane. The data were used to estimate the probable attitudes in steady spins of an airplane in flight and of a model in the free-spinning tunnel. The estimated attitudes of steady spin were compared with attitudes measured in flight and in the spinning tunnel. The results indicate that corrections for certain scale and tunnel effects are necessary to estimate full-scale spinning attitudes from model results.