From Summary: "A method is presented for the rapid calculation of the pressure distribution over an airfoil section when the normal-force distribution and the pressure distribution over the "base profile" (i.e., the profile of the same airfoil were the camber line straight and the resulting airfoil at zero angle of attack) are known. This note is intended as a supplement to N.A.C.A. Report Nos. 631 and 634 wherein methods are presented for the calculation of the normal-force distribution over plain and flapped airfoils, respectively, but not of the pressures on the individual surfaces. Base-profile pressure-coefficient distributions for the usual N.A.C.A. family of airfoils, which are also suitable for several other commonly employed airfoils, are included in tabular form. With these tabulated base-profile pressures and the computed normal-force distributions, pressure distributions adequate for most engineering purposes can be obtained."

From Summary: "Optimum proportions of tapered wings were investigated by a method that involved a comparison of wings designed to be aerodynamically equal. The conditions of aerodynamic equality were equality in stalling speed, in induced drag at a low speed, and in the total drag at cruising speed. After the wings were adjusted to aerodynamic equivalence, the weights of the wings were calculated as a convenient method of indicating the optimum wing. The aerodynamic characteristics were calculated from wing theory and test data for the airfoil sections. Various combinations of washout, camber increase in the airfoil sections from the center to the tips, and sharp leading edges at the center were used to bring about the desired equivalence of maximum lift and center-stalling characteristics. In the calculation of the weights of the wings, a simple type of spar structure was assumed that permitted an integration across the span to determine the web and the flange weights. The covering and the remaining weight were taken in proportion to the wing area. The total weights showed the wings with camber and washout to have the lowest weights and indicated the minimum for wings with a taper ratio between 1/2 and 1/3."

From Summary: "The problem of developing a windshield for aircraft which will withstand the effect of bird impacts during flight is a difficult one, as an estimate of the striking energy will indicate. If the average speed of the airplane is considered to be about 200 miles per hour and that of the bird about 70 miles per hour, the speed of the bird relative to the airplane may be as great as 400 feet per second. If a 4-pound bird is involved, a maximum impact energy of approximately 10,000 foot-pounds must be dissipated. To obtain this energy in a drop test in the Washington Monument, it would be necessary to drop a 20-pound weight down the 500-foot shaft. For both theoretical and practical reasons, it is necessary to keep the mass and speed more nearly like those to be encountered. However, to get an impact of about 10,000 foot-pounds with a 4-pound falling body, it would be necessary to drop it from a height of approximately one-half mile, neglecting air resistance. These facts will indicate some of the experimental obstacles in the way of simulating bird impacts against aircraft windshields."

The present report deals with six-component measurements in the small tunnel of the DVL on a model of the BFW-M 27b(sub 1), which were made to determine the effect of rolling and yawing on the air forces and moments. The wind was given a spiral motion by means of a rotating screen, the model being suspended in the conventional manner. The lack of accuracy in the measurement of the rolling-yawing moments was very noticeable.

"The flow over the inner halves of the rotor blades on a Kellet YG-1B autogiro was investigated in flight by making camera records of the motion of silk streamers attached to the upper surfaces of the blades. These records were analyzed to determine the boundaries of the region within which the flow over the blade sections was stalled for various tip-speed ratios. For the sake of comparison, corresponding theoretical boundaries were obtained. Both the size of the stalled area and its rate of growth with increasing tip-speed ratio were found to be larger than the theory predicted, although experiment agreed with theory with regard to shape and general location of the stalled area" (p. 1).

Note presenting tests of several forms of fixed wing slot in a large-chord NACA 23012 airfoil in the closed-throat 7- by 10-foot wind tunnel. The airfoil extended completely across the test section so that two-dimensional flow was approximated. The data are presented in the form of tables of important aerodynamic characteristics for each slot tested and as curves of section lift, profile-drag, and pitching-moment coefficients.

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.

Note presenting testing of an N.A.C.A. 23012 rectangular wing with rounded tips in combination with a fuselage of circular cross section at several angles of yaw in the NACA 7- by 10-foot wind tunnel. The model was tested as a high-wing, a midwing, and a low-wing monoplane; for each wing location, tests were made with two amounts of dihedral and with partial-span split flaps. Results regarding the wing and fuselage, fin and fuselage, and wing, fuselage, and fin are provided.

In the present paper, the complete buckling conditions of stiffened plates are being developed for uniform compression. We shall treat plates with one or two longitudinal or transverse stiffeners at any point, discuss the buckling conditions, and evaluate them for different cases.

Note presenting testing of twenty-two models of flying-boat hulls for the purpose of determining the effects on water resistance and spray of 13 variations in the transverse section of the bottom of the forebody and of three variations in the form of the afterbody. Generally, the effect of chine flare on the resistance was small, although the resistance of force with chine flare was generally less than the resistance of the form without chine flare.

This paper restricts itself to radial impellers with cylindrical blades since, as Prasil has shown, the flow about an arbitrarily curved surface of revolution may be reduced to this normal form we have chosen by a relatively simple conformal transformation. This method starts from the simple hypotheses of the older centrifugal impeller theory by first assuming an impeller with an infinite number of blades. How the flow is then modified is then investigated.

"The optimum circulation distribution and hence the maximum theoretical thrust obtainable for contra-vanes fitted behind propellers is markedly dependent on the number of guide vanes. The outer portion of the vanes, even if projecting considerably beyond the edge of the propeller slipstream, still contribute appreciably to this theoretical gain of thrust. But, owing to the always existing friction of the vanes, the limit of the optimum vane length lies at relatively small diameters" (p. 1).

Report presenting flight tests to determine the profile-drag coefficients of a portion of the original wing surface of an all-metal airplane and a portion of the wing made aerodynamically smooth and more nearly fair than the original section. Tests were also carried out to determine the point of transition from laminar to turbulent boundary layer and to determine the velocity distribution along the upper surface of the wing.

"Flight tests were made to determine the profile-drag coefficients of a portion of the original wing surface of an all-metal airplane and of a portion of the wing made aerodynamically smooth and more nearly fair than the original section. The wing section was approximately the NACA 2414.5. The tests were carried out over a range of airplane speeds giving a maximum Reynolds number of 15,000,000. Tests were also carried out to locate the point of transition from laminar to turbulent boundary layer and to determine the velocity distribution along the upper surface of the wing" (p. 483).

From Summary: "An analysis was made to determine the proportions of fins made of aluminum, copper, magnesium, and steel necessary to dissipate maximum quantities of heat for different fin widths, fin weights, and air-flow conditions. The analysis also concerns the determination of the optimum fin proportions when specified limits are placed on the fin dimensions. The calculation of the heat flow in the fins is based on experimentally verified, theoretical equations. The surface heat-transfer coefficients used with this equation were taken from previously reported experiments. In addition to the presentation of fin-design information, this investigation shows that optimum fin dimensions are inappreciably affected by the differences in air flow that are obtained with different air-flow arrangements or by small changes in the length of the air-flow path."

From Summary: "Wind-tunnel tests are reported of six 3-blade 10-foot propellers operated in front of a liquid-cooled engine nacelle. The propellers were identical except for blade airfoil sections, which were: Clark y, R.A.F. 6, NACA 4400, NACA 2400-34, NACA 2rsub200, and NACA 6400. The range of blade angles investigated extended for 15 degrees to 40 degrees for all propellers except the Clark y, for which it extended to 45 degrees. The results showed that the range in maximum efficiency between the highest and lowest values was about 3 percent. The highest efficiencies were for the low-camber sections."

"Wind-tunnel tests were conducted on a model wing-nacelle combination to determine the practicability of cooling radial engines by forcing the cooling air into wing-duct entrances located in the propeller slipstream, passing the air through the engine baffles from rear to front, and ejecting the air through an annular slot near the front of the nacelle. The drag of the cowlings tested was definitely less than for the conventional N.A.C.A. cowling, and the pressure available at low air speed corresponding to operation on the ground and at low flying speeds was apparently sufficient for cooling most present-day radial engines" (p. 1).

"This paper is one of several dealing with methods intended to reduce the drag of present-day radial engine installations and improve the cooling at zero and low air speeds. The present paper describes model wind-tunnel tests of blowers of three designs tested in conjunction with a wing-nacelle combination. The principle of operation involved consists of drawing cooling air into ducts located in the wing root at the point of maximum slipstream velocity, passing the air through the engine baffles from rear to front, and exhausting the air through an annular slot located between the propeller and the engine with the aid of a blower mounted on the spinner" (p. 1).

"Two cowling systems intended to reduce the drag and improve the low-speed cooling characteristics of conventional radial engine cowlings were tested in model form to determine the practicability of the methods. One cowling included a blower mounted on the rear face of a large propeller spinner which drew cooling air in through side entrance ducts located behind the equivalent engine orifice plate. The air was passed through the equivalent engine orifice plate from rear to front and out through a slot between the spinner and the engine plate" (p. 1).

Measurements were made of the cooling on the fronts of model cylinders in a conventional cowling for cooling in both the ground and the cruising conditions. The mechanisms of front and rear cooling are essentially different. Cooling on the rear baffled part of the cylinders continually increases with increasing fin width. For the front of the cylinder, an optimum fin width was found to exist beyond which an increase in width reduced the heat transfer. The heat transfer coefficient on the front of the cylinders was larger on the side of the cylinder facing the propeller swirl than on the opposite side. This effect became more pronounced as the fin width was increased. These results are introductory to the study of front cooling and show the general effect of several test parameters.

"The fundamental principles of fluid flow, pressure losses, and heat transfer have been presented and analyzed for the case of a smooth tube with fully developed turbulent flow. These equations apply to tubes with large length-diameter ratios where the flow is at a high Reynolds Number. The error introduced by using these equations increases as the magnitude of the tube length and the air-flow Reynolds Number approaches the values encountered in modern radiator designs" (p. 1).

"When an airplane is operating at high altitude, it is necessary to use a supercharger to maintain ground pressure at the carburetor inlet. This maintenance and high intake-manifold pressure tends to keep the power output of the engine at ground-level value. The air, being compressed by the supercharger, however, is heated by adiabatic compression and friction to a temperature that seriously affect the performance of the engine. It is thus necessary to use an intercooler to reduce the temperature of the air between the supercharger outlet and the carburetor inlet" (p. 1).

"The strength of representative types of flush-riveted joints has been determined by testing 865 single-shearing, double-shearing, and tensile specimens representing 7 types of rivet and 18 types of joint. The results, presented in graphic form, show the stress at failure, type of failure, and d/t ratio. In general, dimpled joints were appreciably stronger than countersunk or protruding-head joints, but their strength was greatly influenced by constructional details" (p. 467).

Direct measurements were undertaken at the Aeronautics Laboratory in Turin of the aerodynamic actions on an oscillating wing. The tests conducted had as their essential object the examination of the operation of apparatus designed for this measurement. The values experimentally obtained for the aerodynamic coefficients are in good agreement with the theory of oscillatory motion of the wing of finite span and show clear deviation from the values obtained by theory of plane motion.