ALBERT

Beschreibung: This report deals with the development of a method which gives a lucid and convenient solution of the flow conditions in the vicinity of a common, thick airfoil section wherein the thickness of the profile is taken into account. The method consists in making the airfoil the streamline in a parallel flow by disposing on its mean line certain source and vortex distributions the fields of which are superposed on the parallel flow. These distributions of singularities are secured for the generalized Karman-Trefftz profile by means of conformal transformation from the flow about a circle. Five different distribution functions are afforded for the density of superposition, which combine in a specified manner to the necessary distributions of singularity and represent a generalized Karman-Trefftz profile in parallel flow. For these profiles the speed for each of the five distributions is then computed independently of the angle of attack.

Beschreibung: Wing pressure distribution diagrams for several angles of attack and flap deflections of 0 degrees, 20 degrees, and 40 degrees are presented. The normal force coefficients agree with lift coefficients obtained in previous test of the same model, except for the maximum lifts with flap deflection. Pressure distribution measurements were made at Reynolds Number of about 6,000,000.

Beschreibung: At the request of the Army Air Forces, Materiel Command, tests were made in the two-Dimensional low-turbulence tunnel of a mode l submitted by Bell Aircraft Corporation as the tip section of the wing of the xp-63 airplane. " The model was a NACA 66, 2x-2l6 a = 0.6 airfoil section of 50-inch chord and made of dural according to standard sheet-metal practice. The model was fitted for an internal-balance aileron with a tab and a Frise type aileron. Several alterations of the skirts and balance on the internal- balance aileron were tested to obtain the various aerodynamic characteristics of the aileron.

Beschreibung: In many studies, especially of nonstationary flight motion, it is necessary to determine the angular velocities at which the airplane rotates about its various axes. The three-component recorder is designed to serve this purpose. If the angular velocity for one flight attitude is known, other important quantities can be derived from its time rate of change, such as the angular acceleration by differentiations, or - by integration - the angles of position of the airplane - that is, the angles formed by the airplane axes with the axis direction presented at the instant of the beginning of the motion that is to be investigated.

Beschreibung: The wall interference is obtained for a wind tunnel of elliptic section for the two cases of closed and open working sections. The approximate and exact methods used gave results in practically good agreement. Corresponding to the result given by Glauert for the case of the closed rectangular section, the interference is found to be a minimum for a ratio of minor to major axis of 1:square root of 6 This, however, is true only for the case where the span of the airfoil is small in comparison with the width of the tunnel. For a longer airfoil the favorable ellipse is flatter. In the case of the open working section the circular shape gives the minimum interference.

Beschreibung: The present report deals with a number of problems, particularly with the interaction of the fuselage with the wing and tail, on the basis of simple calculating method's derived from greatly idealized concepts. For the fuselage alone it affords, in variance with potential theory, a certain frictional lift in yawed flow, which, similar to the lift of a wing of small aspect ratio, is no longer linearly related to the angle of attack. Nevertheless there exists for this frictional lift something like a neutral stability point the position of which on oblong fuselages appears to be associated with the lift increase of the fuselage in proximity to the zero lift, according to the present experiments. The Pitching moments of the fuselage can be determined with comparatively great reliability so far as the flow conditions in the neighborhood of the axis of the fuselage can be approximated if the fuselage were absent, which, in general, is not very difficult. For the unstable contribution of the fuselage to the static longitudinal stability of the airplane it affords comparatively simple formulas, the evaluation of which offers little difficulty. On the engine nacelles there is, in addition a very substantial wing moment contribution induced by the nonuniform distribution of the transverse displacement flow of the nacelle along the wing chord; this also can be represented by a simple formula. A check on a large number of dissimilar aircraft types regarding the unstable fuselage and nacelle moments disclosed an agreement with the wind-tunnel tests, which should be sufficient for practical requirements. The errors remained throughout within the scope of instrumental accuracy.

Beschreibung: This report gives the results of tests on a rectangular wing model with a 20% full spun split flap, conducted on the whirling arm at the Daniel Guggenheim Airship Institute in Akron, Ohio. The effect of a ground board on the lift and pitching moment was measured. The ground board consisted of an inclined ramp rising up in the test channel to a level floor extending for some distance parallel to the model path. The path of the wing model with respect to the ground board accordingly represented with comparative exactness an airplane coming in for a landing. The ground clearances over the level portion of the board varied from 0 6 to 1,6 chord lengths. Results are given in the standard dimensionless coefficients plotted versus angle of attack for a particular ground clearance. The effect of the ground board is to increase the lift coefficient for a given angle of attack all the way up the stall. The magnitude of the increase varies both with the ground clearance and the angle of attack. The effect on the pitching moment coefficient is not so readily apparent due to experimental difficulties but, in general, the diving moment increases over the ground board. This effect is apparent principally at the high angles of attack. An exception to this effect occurs with flaps deflected at the lowest ground clearance (0.6 chords). Here the diving moment decreases over the ground board.

Beschreibung: At the present time there is considerable demand for improvement in the aerodynamic characteristics of cowlings for radial air-cooled aircraft engines. During the past year, numerous cowling arrangements have been investigated in various departments of the NACA laboratory. Although a few full-scale investigation have been carried out, most of the studies have been preliminary in nature and have been confined to the investigation of model arrangement in wind tunnels. Because of the existing national emergency it appears advisable to release immediately to the aircraft industry the information available on the more promising of the arrangements that have been studied. An investigation having as its aim the improvement in performance and flying qualities of single-engine air-cooled military pursuit airplanes is being conducted in the NACA 10-foot pressure wind tunnel. As a part of that investigation, studies have been made of the relative merits of a conventional NACA open-nose cowling arrangement and of a less conventional but better streamline NACA high-speed cowling arrangement in which the cooling air enters the cowling through an opening ahead of the propeller, passes internally through an element of the cowling which rotates with the propeller, and thence past the engine cylinders to the exit at the rear of the engine. These investigations indicate that at airplane speeds of around 400 miles per hour there is not a great deal to be gained in high-speed performance through the application of the latter cowling arrangement, but at speeds in excess of about 450 miles per hour a very appreciable gain is indicated. Present indications are that improved engine cooling can be obtained throughout the speed range as well as ground cooling through the use of the high-speed cowling. This paper summarizes the results obtained from wind-tunnel tests of models of the two cowling arrangements.

Beschreibung: In this report a method is presented for the calculation of the profile drag of airfoil sections. The method requlres only a knowledge of the theoretical velocity distribution and can be applied readily once this dlstribution is ascertained. Comparison of calculated and experimental drag characteristics for several airfoils shows a satisfactory agreement. Sample calculatlons are included.

Beschreibung: Investigations of the pressure distribution, the profile drag, and the location of transition for a 30-inch-chord 25-percent-thick N.A,C.A. 45-125 airfoil were made in the N.A.C.A 8-foot high-speed wind tunnel for the purpose of aiding in the development of a thick wing for high-speed airplanes. The tests were made at a lift coefficient of 0.1 for Reynolds Numbers from 1,750,000 to 8,690,000, corresponding to speeds from 80 to 440 miles per hour at 59 F. The effect on the profile drag of fixing the transition point was also investigated. The effect of compressibility on the rate of increase of pressure coefficients was found to be greater than that predicted by a simplified theoretical expression for thin wings. The results indicated that, for a lift coefficient of 0.1, the critical speed of the N.A.C,A. 45-125 airfoil was about 460 miles per hour at 59 F,. The value of the profile-drag coefficient at a Reynolds Number of 4,500,000 was 0.0058, or about half as large as the value for the N.A,C,A. 0025 airfoil. The increase in the profile-drag coefficient for a given movement of the transition point was about three times as large as the corresponding increase for the N.A.C,A. 0012 airfoil. Transition determinations indicated that, for Reynolds Numbers up to ?,000,000, laminar boundary 1ayers were maintained over approximately 40 percent of the upper and the lower surfaces of the airfoil.