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This is the thirteenth report on a series of systematic tests comparing lateral control devices with particular reference to their effectiveness at high angles of attack. The present wind tunnel tests were made to determine the most feasible locations for lateral control surfaces mounted externally to a rectangular Clark y wing.
This report deals with the results of tests on struts conducted at Washington Navy Yard. Two models of the modified Navy strut, no. 1, were tested in the 8 by 8 foot wind tunnel. The tests were made to determine the total resistance end effect and the pressure distribution at various wind-tunnel speeds with the length of the strut transverse to the current. Only the measurements made at zero pitch and yaw are given in this report.
This report covers the twelfth of a series of tests conducted to compare different lateral control devices with particular reference to their effectiveness at high angles of attack. The present wind tunnel tests were made with two sizes of upper-surface ailerons on rectangular Clark Y wing models equipped with full span split flaps. The tests showed the effect of the upper-surface ailerons and of the split flaps on the general performance characteristics of the wings, and on the lateral controllability and stability characteristics. The results are compared with those for plain wings with ordinary ailerons of similar sizes.
Three model wings, two with typical slotted ailerons and one with typical frise ailerons, have been tested as part of a general investigation on lateral control devices with particular reference to their effectiveness at high angles of attack, in the 7 by 10 foot wind tunnel of the National Advisory Committee for Aeronautics. Force tests, free-autorotation tests, and forced-rotation tests were made which show the effect of the various ailerons on the general performance of the wing, on the lateral controllability, and on the lateral stability, in general, rolling control at 20 degree angle of attack to plain ailerons of the same size. The adverse yawing moments obtained with the slotted and frise ailerons were, in most cases, slightly smaller than those obtained with plain ailerons of the same size and deflection. However, this improvement was small as compared to the improvement obtainable by the use of suitable differential movements with any of the ailerons, including the plain.
An investigation was conducted to furnish data on the effect of tab balance on tab and control-surface characteristics. The airfoil tested had a modified NACA 65(1)-012 contour with a plain flap having a chord equal to 25 percent of the wing chord and with a tab having a chord equal to 25 percent of the flap chord and having several nose shapes and overhang lengths. The results of the investigation indicated that, in general, tab balance affected tab hinge-moment characteristics in much the same manner that flap balance affects flap hinge-moment characteristics. A moderate amount of tab balance did not seem to have any adverse effect on flap hinge-moment characteristics.
A wind-tunnel study has been made to determine the possibility of developing spoilers suitable for providing the lateral control for airplanes in place of the usual ailerons. The first tests were made on a model wing with a fixed tip slot, but when it was found that the effectiveness of the spoilers did not depend to any great extent on the slot, tests were made on a plain wing also. In both cases certain spoiler positions were found which were free from the usual adverse rolling moments with small deflections. Five different forms of spoiler were tested, the best ones being simple plates, either straight or slightly curved to fit the contour of the airfoil when not deflected. Sufficient rolling moment can probably be obtained from spoilers of reasonable size to provide satisfactory lateral control for certain types of airplanes.
This report covers the fifth of a series of systematic investigations in which lateral control devices are compared with particular reference to their effectiveness at high angles of attack. The present report deals with tests of spoilers and ordinary ailerons on rectangular Clark y wing models. In an effort to obtain satisfactory control throughout the entire angle-of-attack range that can be maintained in flight, various spoilers were tested in combination with two sizes of previously tested ordinary ailerons - one of average proportions and the other short and wide. In addition, one large spoiler was tested alone. It was found that when ailerons and spoilers are used together the full effect of both is not obtained if the spoilers are located directly in front of the ailerons. With the proper combination of spoiler and aileron, however, it is possible to obtain satisfactory rolling control up to high angles of attack (15 degrees to 20 degrees), together with favorable yawing moments and small control forces. A moderate amount of rolling control with favorable yawing moments and small control forces was obtained with the large spoiler alone.
This report covers the sixth of a series of investigations in which various lateral control devices are compared with particular reference to their effectiveness at high angles of attack. The present report deals with flap-type ailerons hinged about axes having an angle with respect to the leading and trailing edges of the wing. Tests were made on four different skewed ailerons, including two different angles of skew and two sizes of ailerons. At the high angles of attack, all the skewed ailerons tested were slightly inferior with respect to rolling and yawing moments to straight ailerons having the same span and average chord. Computations indicate that the skewed ailerons are also inferior with respect to hinge moments.
The tests described in this report furnished data on the actual aerodynamic forces, and the resulting stresses and bending moments in the hull of the U. S. S. "Los Angeles" during as severe still-air maneuvers as the airship would normally be subjected to, and in straight flight during as rough air as is likely to occur in service, short of squall or storm conditions. The maximum stresses were found to be within the limits provided for in accepted practice in airship design. Normal flight in rough air was shown to produce forces and stresses about twice as great as the most severe still-air maneuvers. No light was thrown upon the forces which might occur in extreme or exceptional conditions, such as the storm which destroyed the "Shenandoah". The transverse aerodynamic forces on the hull proper were found to be small and irregular. Owing to the necessity of conserving helium, it was impossible to fly the airship in a condition of large excess of buoyancy or weight in order to determine the air pressure distribution at a fixed angle of pitch. However, there is every reason to believe that in that condition the forces on the actual airship are as close to the wind-tunnel results as can be determined by present type of pressure measuring apparatus. It is considered that most important data obtained are the coefficients of tail-surface forces and hull-bending moments. These are tabulated in this report.
Some preliminary results of full scale wind tunnel testing to determine the best means of reducing the tail buffeting and wing-fuselage interference of a low-wing monoplane are given. Data indicating the effects of an engine cowling, fillets, auxiliary airfoils of short span, reflexes trailing edge, propeller slipstream, and various combinations of these features are included. The best all-round results were obtained by the use of fillets together with the National Advisory Committee for Aeronautics (NACA) cowling. This combination reduced the tail buffeting oscillations to one-fourth of their original amplitudes, increased the maximum lift 11 percent, decreased the minimum drag 9 percent, and increased the maximum ratio of lift to drag 19 percent.
This paper gives some quantitative measurements of wind tunnel turbulence and its effect on the air resistance of spheres and airship models, measurements made possible by the hot wire anemometer and associated apparatus in its original form was described in Technical Report no. 320 and some modifications are presented in an appendix to the present paper. One important result of the investigation is a curve by means of which measurements of the air resistance of spheres can be interpreted to give the turbulence quantitatively. Another is the definite proof that the discrepancies in the results on the N. P. L. Standard airship models are due mainly to differences in the turbulences of the wind tunnels in which the tests were made. An attempt is made to interpret the observed results in terms of the boundary layer theory and for this purpose a brief account is given of the physical bases of this theory and of conceptions that have been obtained by analogy with the laws of flow in pipes.
The theory of wind tunnel boundary influence on the downwash from a wing has been extended to provide more complete corrections for application to airplane test data. The first section of the report gives the corrections of the lifting line for wing positions above or below the tunnel center line; the second section shows the manner in which the induced boundary influence changes with distance aft of the lifting line. Values of the boundary corrections are given for off-center positions of the wing in circular, square, 2:1 rectangular, and 2:1 elliptical tunnels. Aft of the wing the corrections are presented for only the square and the 2:1 rectangular tunnels, but it is believed that these may be applied to jets of circular and 2:1 elliptical cross sections. In all cases results are included for both open and closed tunnels.
This report presents the results of wind tunnel tests to determine the effectiveness of various methods of reducing the high control forces of unbalanced upper-surface ailerons and of unbalanced split flaps.
Report presents the results of wind tunnel tests of the Clark Y, the NACA 23012 and the NACA 23021 airfoils equipped with full-span ordinary flaps and with full-span simple split flaps. The principal object of the tests was to determine the characteristics of the airfoils with ordinary flaps and in addition, to determine the relative merits of the various airfoils when equipped with either ordinary flaps or with simple split flaps. The Clark Y airfoil was tested with 3 widths of ordinary flap, 10, 20, and 30 percent of the airfoil chord. The optimum width of the ordinary and the simple split flap based on the maximum lift attained with the Clark Y airfoil was then tested on each of the other two airfoils.
Report presents an investigation carried out in the NACA 7- by 10-foot wind tunnel of an NACA 23012 airfoil equipped, first, with a full-span NACA 23012 external-airfoil flap having a chord 0.20 of the main airfoil chord and with a full-span aileron with a chord 0.12 of the main airfoil chord on the trailing edge of the main airfoil and equipped second, with a 0.30-chord full-span NACA 23012 external-airfoil flap and a 0.13-chord full-span aileron. The results are arranged in three groups, the first two of which deal with the airfoil characteristics of the two airfoil-flap combinations and with the internal-control characteristics of the airfoil-flap-aileron combinations. The third group of tests deals with several means for balancing ailerons mounted on a special large-chord NACA 23012 external-airfoil flap. The tests included an ordinary aileron, a curtained-nose balance, a frise balance, and a tab.
The interference of the wind-tunnel boundaries on the downwash behind an airfoil has been experimentally investigated and the results have been compared with the available theoretical results for open-throat wind tunnels. As in previous studies, the simplified theoretical treatment that assumes the test section to be an infinite free jet has been shown to be satisfactory at the lifting line. The experimental results, however, show that this assumption may lead to erroneous conclusions regarding the corrections to be applied to the downwash in the region behind the airfoil where the tail surfaces are normally located. The results of a theory based on the more accurate concept of the open-jet wind tunnel as a finite length of free jet provided with a closed exit passage are in good qualitative agreement with the experimental results.
Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of large-chord NACA and 23021, and 23030 airfoils with split flaps 10, 20, 30, and 40 percent of the wing chord to determine the section aerodynamic characteristics of the airfoils as affected by airfoil thickness, flap chord, and flap deflection. The complete section aerodynamic characteristics of all the combinations tested are given in the form of graphs of lift, drag, and pitching-moment coefficients, and certain applications to aerodynamic design are discussed.
Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of a large-chord NACA 23012 airfoil with several arrangements of venetian-blind flaps to determine the aerodynamic section characteristics as affected by the over-all flap chord, the chords of the slats used to form the flap, the slat spacing, the number of slats and the position of the flap with respect to the wing. Complete section data are given in the form of graphs for all the combinations tested.