"Pressure-distribution tests were made on the 32-foot whirling arm of the Daniel Guggenheim Airship Institute of a tapered wing to determine the rolling and yawing moments due to an angular velocity in yaw. The model was tested at 0 degree and 5 degrees pitch, -1 degree and 5 degree yaw, and with a full-span flap deflected 60 degrees. The results are given in the form of span load distributions and in calculated moment coefficients" (p. 1).

Note discussing the lift of a finite wing including the calculations and constants for each step of the lift. This information is compared against the lift of an infinite wing. From Summary: "Unsteady lift function for wings of finite aspect ratio have been calculated by approximate methods involving corrections of the aerodynamic inertia and of the angle of the infinite wing. The starting lift of the finite wing is found to be only slightly less than that of the infinite wing; whereas the final lift may be considerably less. The calculations indicate that the distribution of lift near the start is similar to the final distribution."

"A flight investigation was made of the increase in effectiveness of ailerons that can be obtained by preventing flow of air through the wing at the hinges and of the possibility of reducing the aileron operating force by replacing ailerons having normal open hinge gaps with narrower but equally effective ailerons having sealed hinge gaps. Tests were made with a Fairchild 22 airplane with two sizes of plain unbalanced ailerons, one set having a chord equal to 0.18c, and the other chord equal to 0.09c. The results of the investigation show that improvement of the lateral-control effectiveness is obtained by completely preventing the flow of air through the wing at the hinge axis of conventional ailerons" (p. 1).

An optically recording engine-pressure indicator of simple and rugged construction has been developed for use in high-pressure and high temperature combustion research. This instrument is of the diaphragm type and has a natural frequency of about 10,000 cycles per second.

Three rectangular monoplane wings having rounded tips were tested on the N.A.C.A. spinning balance in the 5-foot vertical wind tunnel. The airfoil sections used were the N.A.C.A. 0009, 23018, and 6718. The aerodynamic characteristics of the models and a prediction of the angles of sideslip for steady spins are given. There is included an estimate of the yawing moment that must be furnished by parts of the airplane to balance the inertia couples and wing yawing moments for spinning equilibrium. The predicted angles of sideslip and yawing moments required for spinning equilibrium for a Clark Y wing with the same form are included for comparison.

"An investigation was conducted in the NACA 20-foot wind tunnel of the ramming effect of three general types of carburetor intake rams for radial engines, namely, the internal constant area type, the external constant area type, and the external expanding type. The rams were installed on a radial air- cooled engine nacelle, and tests were made with and without the propeller operating. The results indicated that the external types having entrances near the front of the engine cowling gave the greatest ramming effect" (p. 1).

"The hydrodynamic characteristics of a model of the hull of the Short Calcutta (N.A.C.A. Model 47) are presented in non-dimensional form. This model represents one of a series of hulls of successful foreign and domestic flying boats the characteristics of which are being obtained under similar test conditions in the N.A.C.A. tank. The take-off distance and time for a flying boat having the hull of the Calcutta are compared at two values of the gross load with the corresponding distances and times for the same flying boat having hulls of two representative American types, the Sikorsky S-40 and the N.A.C.A. 11-A" (p. 1).

"Extensive strain-gage measurements were made chiefly on the tension side of five small rectangular box beams constructed of sheet duralumin. The main conclusion was that within the test range the tension cover may be considered as being fully effective but that at any given point on the beam there may be unaccountable variations of 5 percent from the calculated stresses on the thicker sheets tested (0.044 and 0.023 inch) and of 10 percent or more on the thinnest sheet tested (0.014 inch)" (p. 1).

"The effect of tapered ailerons on linearly tapered wings is theoretically determined. Four different aileron spans are considered for each of three wing aspect ratios and each of four wing taper ratios. The change in lift on one half of the wing, the rolling moment, the additional induced drag, and the yawing moment, due to aileron deflection, are represented by non dimensional coefficients" (p. 1).

Measurements were made to determine the relative maximum forces a pilot can exert on the controls of an airplane with the view of obtaining systematic data upon which to base the location of controls within the cockpit and the design of the control surfaces. A cockpit model of generous proportions, capable of being rotated to any attitude, was built with the location of the control stick and rudder pedals adjustable over a wide range of positions with respect to the seat. Besides measurements of maximum forces obtainable with various control locations and with the pilot in several attitudes, estimates of forces within the range normally encountered in flight were made to gain an indication of the accuracy of estimating control forces.

"The proper landing of airplanes equipped with flaps, although probably no more difficult than landing without them, requires a different technique. The effects of flaps on the aerodynamics characteristics of a wing are given and, with the aid of figures and diagrams, a detailed comparison of the glide and landing of an airplane with and without flaps is made. The dangers attending improper execution and the importance of such factors as air speed fuselage attitude, glide-path angle, and control manipulation, upon all of which a pilot bases his judgement, are emphasized" (p. 1).

From Summary: "The results of tank tests of three models of flying-boat hulls of the pointed-step type with different angles of dead rise are given in charts and are compared with results from tests of more conventional hulls. Increasing the angle of dead rise from 15 to 25 degrees: had little effect on the hump resistance; increased the resistance throughout the planning range; increased the best trim angle; reduced the maximum positive trimming moment required to obtain best trim angle; and had but a slight effect on the spray characteristics. For approximately the same angles of dead rise the resistance of the pointed-step hulls were considerably lower at high speeds than those of the more conventional hulls."

"Preliminary tests have been made for the purpose of obtaining a flap arrangement suitable for direct and immediate control of the steepness of the glide path of an airplane, a use for which present flaps are not satisfactory. An attempt has been made to develop a flap giving a reasonably high maximum lift coefficient with relatively low deflection and maintaining this value of the maximum lift coefficient with a large increase of deflection, the increase in deflection being accompanied by a large increase in drag. An arrangement was found that gave a maximum lift coefficient of approximately 1.90 for all flap deflections between 25 and 80 degrees, within which range the drag of the wing increased regularly to a large value" (p. 1).

"The ignition lag of a fuel oil in the combustion chamber of a high speed compression-ignition engine was measured by three different methods. The start of injection of the fuel as observed with a Stoborama was taken as the start of the period of ignition lag in all cases. The end of the period of ignition lag was determined by observation of the appearance of incandescence in the combustion chamber, by inspection of a pressure-time card for evidence of pressure rise, and by analysis of the indicator card for evidence of the combustion of a small but definite quantity of fuel" (p. 1).

From Summary: "Data obtained at the N.A.C.A. tank from tests on the models of three flying-boat hulls - N.A.C.A. models 11-A, 16, and 22 - are used to demonstrate the effect of trim angle on water resistance. A specific example is taken, and data from Model 11-A are used to show that the trim angle giving the minimum water resistance will give minimum total air-plus-water resistance. Total-resistance curves for best trimmed angles and other angles are compared for the same example."

"This paper contains a series of calculations showing how the performance of controllable propellers may be derived from data on fixed-pitch propellers given in N.A.C.A. Technical Report No. 350, or from similar data. Sample calculations are given which compare the performance of airplanes with fixed-pitch and with controllable propellers. The gain in performance with controllable propellers is shown to be largely due to the increased power available, rather than to an increase in efficiency" (p. 1).

Three tapered airfoils based on the N.A.C.A. 2200, the N.A.C.A.-M6, and the Clark Y sections were tested in the variable-density wind tunnel at a Reynolds Number of approximately 3,100,000. The models, which were of aspect ratio 6, had constant core center sections and rounded tips, and tapered in thickness from 18 percent at the roots to 9 percent at the tips. The aerodynamic characteristics are given by the usual dimensionless coefficients plotted for both positive and negative angles of attack and by effective profile-drag coefficients plotted against lift coefficients.

An investigation of the possibility of using thrust augmented jets as prime movers was carried out. The augmentation was to be effected by allowing the jet to mix with the surrounding air in the presence of bodies which deflect the air set in motion by the jet. Six augmentation schemes were tested experimentally with compressed air at room temperature at jet speeds up to 1240 feet per second.

Ailerons described in references 1, 2, and 3, are summarized in the form of empirical equations which relate the aileron dimensions and displacements to the rolling, yawing, and hinge moments for pitch angles of 0 and 12 degrees, corresponding to angles of attack of the wings of 4 and 16 degrees, respectively.

Tests were made with ordinary ailerons and different sizes of spoilers on rectangular Clark Y wing models with Handley Page tip and full span slots. The tests showed the effect of the control devices on the general performance of the wings as well as on the lateral control and lateral stability characteristics.

This report presents charts which reduce the stress analysis of circular and elliptic rings of uniform cross section subjected to balanced systems of concentrated loads from a statically indeterminate problem to a statically determinate one. To demonstrate the use of the charts in the stress analysis of elliptic rings, an illustrative problem is included.

"This note is the fifth of a series covering an investigation of a number of related airfoils. It presents the results obtained from tests of a group of six low-cambered airfoils in the variable-density wind tunnel. The mean camber lines are identical for the six airfoils and are of such a form that the maximum mean camber is 2 per cent of the chord and is at a position 0.4 of the chord behind the loading edge. The airfoils differ in thickness only, the maximum-thickness/chord ratios being 0.06, 0.09, 0.12, 0.15, 0.18, and 0.21" (p. 1).

This report deals with the effect of small variations in ordinates specified by different laboratories for the airfoil section. This study was made in connection with a more general investigation of the effect of small irregularities of the airfoil surface on the aerodynamic characteristics of an airfoil. These tests show that small changes in airfoil contours, resulting from variations in the specified ordinates, have a sufficiently large effect upon the airfoil characteristics to justify the taking of great care in the specification of ordinates for the construction of models.

From Summary: "The results of flight tests conducted by the NACA to determine the lift and drag characteristics of a full-scale airplane are given herein. A Fairchild FC-2W2 cabin monoplane having a Gottingen 387 wing section was used for the tests. The maximum lift coefficient for the airplane is compared with that obtained for the Gottingen 387 airfoil in recent tests in the Variable Density Tunnel. The maximum lift coefficient for the airplane was found to be 1.50 and that for the airfoil 1.56. Although the flight tests were confined chiefly to glides with the propeller locked horizontally, data obtained with the propeller operating at zero thrust for a few angles of attack are also included."