"The 'Farnboro' electric indicator was tested as received from the manufacturers, and modifications made to the instrument to improve its operation. The original design of disk valve was altered so as to reduce the mass, travel, and seat area. Changes were made to the recording mechanism, which included a new method of locating the top center position on the record. The effect of friction on the motion of the pointer while taking motoring and power cards was eliminated by providing a means of putting pressure lines on the record" (p. 1).
"Charts are presented for determining the performance of airplanes having variable-pitch propellers, the pitch of which is assumed to be adjusted to maintain constant speed for all rates of flight. The charts are based on the general performance equations developed by Oswald in reference 1, and are used in a similar manner. Examples applying the charts to airplanes having both supercharged and unsupercharged engines are included" (p. 1).
The results of a research program to obtain design data on the strength of open-channel aluminum-alloy sections subjected to combined column and beam action. The results of the tests of about 70 specimens were graphed for stresses due to axial load and stresses due to bending loading as functions of length to radius of gyration of the specimens. From these graphs a design chart was derived that is suitable for ready use.
Model No. 11-A was designed as an improvement over N.A.C.A. Model No. 11, a complete test of which is described in N.A.C.A. Technical Note No. 464. In contrast with the longitudinal upward curvature in the planing bottom forward of the main step on Model 11-A was made as flat as practicable. Otherwise, the two models have very nearly the same form. The results of towing tests made on Model 11-A in the N.A.C.A. tank over a wide range of speed, load on the water, and trim angle are presented, both as original test data and as non dimensional coefficients. A comparison is made with similar results from the test of Model No. 11. The practical significance of the improvement obtained is demonstrated by applying the data from the new form to the illustrative design problem use in the note on Model NO. 11.
"A model of a 2-step flying-boat hull, of the type generally used in England, was tested according to the complete method described in the N.A.C.A. Technical Note No. 464. The lines of this model were taken from offsets given by Mr. William Munro in Flight, May 29, 1931. The data cover the range of loads, speeds, and trim angles that may be of use in applying the hull form to the design of any seaplane. The results are reduced to nondimensional form to aid application to design problems and facilitate comparison with the performance of other hulls" (p. 1).
"This note presents the results of complete tank test of N.A.C.A. Models 22-A and 35, two flying-boat hulls of the deep pointed-step type with low dead rise. Model 22-A is a form derived by modification of Model 22, the test results of which are given in N.A.C.A. Technical Note No. 488. Model 35 is a form of the same type but has a higher length-beam ratio than either Model 22 or 22-A. Take-off examples are worked out using data from these tests and a previous test of a conventional model applied to an arbitrary set of design specifications for a 15,000-pound flying boat" (p. 1).
In this paper there is presented a deflection formula for single-span beams of constant section subjected to combined axial and transverse loads of the types commonly encountered in airplane design. The form of the equation is obtainable by dimensional analysis. Tables and curves of the non dimensional coefficients are appended to facilitate the use of the formula. The equation is applied to the determination of the spring constant of a beam. Tables and curves are presented to show the variation of the spring constant with changes in the axial load and position along the beam.
Note presenting a discrepancy between the predicted and actual point of transition from laminar to turbulent boundary layer that had been found. This effect may be due to the comparatively small radius of curvature of the upper surface of the wing.
"In the flow over the upper surface of a wing, a discrepancy between the predicted and actual point of transition from laminar to turbulent boundary layer was found. This effect may be due to the comparatively small radius of curvature of the upper surface of the wing. Tests were undertaken to investigate this effect. As far as the authors know, the present investigation is the first to show that curvature has a pronounced effect on the transition of the boundary layer from the laminar to the turbulent state" (p. 1).
"Aerodynamic force tests were made in the N.A.C.A. 7 by 10 foot wind tunnel on a model Clark Y wing with a 20 percent chord split flap deflected 60 degrees downward. The tests were made to determine the effect of partial-span split flaps, located at various positions along the wing span on the aerodynamic characteristics of the wing-and-flap combination. The different lengths and locations of the flaps were obtained by cutting off portions of a full-span flap, first from the tips and then from the center. The results are given in the form of curves of lift, drag, and center of pressure" (p. 1).
NACA model 11-C was tested according to the general method with the angle of afterbody keel set at five different angles from 2-1/2 degrees to 9 degrees, but without changing other features of the hull. The results of the tests are expressed in curves of test data and of non-dimensional coefficients. At the depth of step used in the tests, 3.3 percent beam, the smaller angles of afterbody keel give greater load-resistance ratios at the hump speed and smaller at high speed than the larger angles of afterbody keel. Comparisons are made of the load-resistance ratios at several other points in the speed range. The effect of variation of the angle of afterbody keel upon the take-off performance of a hypothetical flying boat of 15,000 pounds gross weight having a hull of model 11-C lines is calculated, and the calculations show that the craft with the largest of the angles of afterbody keel tested, 9 degrees, takes off in the least time and distance.
"The investigation was made to determine the effects of full-span and of partial-span split flaps on the aerodynamic characteristics of a tapered wing. Aerodynamic force tests were made in the N.A.C.A. 7 by 10 foot wind tunnel on a highly tapered Clark Y wing equipped with various split flaps. Two sizes of tapered-chord flaps were tested as full-span flaps, and a narrow tapered-chord flap was tested as a partial-span flap by cutting off portions first from the tip and then from the center" (p. 1).
An investigation was made to determine the aerodynamic characteristics of tapered and rectangular wings with partial-span plain flaps. Two Clark Y airfoils equipped with center section and with tip-section flaps were tested. The results showed that the aerodynamic characteristics of partial-span plain flaps were, in general, similar to those of split flaps of the same span, but that the lift and the drag were less for the wing with plain flaps than for the wing with split flaps of comparable size. For the rectangular wing with center-section plain flaps, the maximum lift and the lift-drag ratio at maximum lift were greater and the drag at maximum lift was less than for the wing with tip-section plain flaps of the same size. The maximum lift of the tapered wing varied in the same manner as that of the rectangular wing but the drag and the lift-drag-ratio relationship were opposite.
Report presenting an investigation of the flow of ice-inhibiting fluids over the blade surfaces of a 12.5-foot-daimeter propeller in flight by discharging dyed fluids at various stations along the leading edges of the blades. The effects on the distribution of varying the fluid composition, the blade-surface roughness, and the orifice design were also observed.
Report presenting an investigation of 24 wing-tail combinations with the weight moved from the center of gravity toward the wing tips so that the distribution of mass along the wings was increased. Results regarding the effects of wings, effects of tail arrangement, effects of control setting, relationships between spin characteristics, and comparison with results for basic loading are provided.
From Summary: "Results are presented of the drag tests of six bodies of revolution with systematically varying shapes and with a fineness ratio of 5. The forms were derived from source-sink distributions, and formulas are presented for the calculation of the pressure distribution of the forms. The tests were made in the N.A.C.A. variable-density tunnel over a range of values of Reynolds number from about 1,500,000 to 25,000,000. The results show that the bodies with the sharper noses and tails have the lowest drag coefficients, even when the drag coefficients are based on the two-thirds power of the volume. The data shows the most important single characteristic of the body form to be the tail angle, which must be fine to obtain low drag."
A general test was made in the N.A.C.A. tank of a 1/12-size model of the hull of the British Singapore IIC flying boat loaned by the Director of Research, British Air Ministry. The results are given in charts and are compared with the results of tests of a model of an American flying-boat hull, the Sikorsky S-40. The Singapore hull has a greater hump resistance but a much lower high-speed resistance than the S-40.
The gyroscopic instruments commonly used in instrument flying in the United States are the turn indicator, the directional gyro, the gyromagnetic compass, the gyroscopic horizon, and the automatic pilot. These instruments are described. Performance data and the method of testing in the laboratory are given for the turn indicator, the directional gyro, and the gyroscopic horizon. Apparatus for driving the instruments is discussed.
"A method of calculating the stresses and torsional stiffness of thin shells with interior webs is summarized. Comparisons between experimental and calculated results are given for 3 duralumin beams, 5 stainless steel beams and 2 duralumin wings. It is concluded that if the theoretical stiffness is multiplied by a correction factor of 0.9, experimental values may be expected to check calculated values within about 10 percent" (p. 1).
Technical note presenting the design and use of a spring-loaded tautness meter that can be used in both horizontal and vertical positions to compare the tautness of various panels. Results of tests made on the fabric coverings of various airplanes are reported and discussed.
"This paper describes an investigation on airplane landing speeds which was made to determine the applicability of accepted aerodynamic theory to the prediction of this particular performance characteristic. The experimental work consisted in measuring the landing speed of several monoplanes by a new photographic method. The results of these tests supplemented by available information regarding biplanes were compared with predictions made with basic aerodynamic theory" (p. 1).
"This is a resume of the work of the Bureau of Standards on a cotton parachute cloth for use as a substitute for silk in the event of an emergency curtailing the supply. Cotton yarn of high strength in proportion to its weight and otherwise specially suitable for parachute cloth was developed. Cloth woven from this yarn in the bureau mill was equal or superior to parachute silk in strength and tear resistance, met the requirements with respect to air permeability, and weighed only a few tenths of an ounce per square yard more than the silk cloth" (p. 1).
"The effect of restraint in pitching introduced by an automatic pilot upon the longitudinal stability of an airplane has been studied. Customary simplifying assumptions have been made in setting down the equations of motion, and the results of computations based on the simplified equations are presented to show the effect of an automatic pilot installed in an airplane of known dimensions and characteristics. The equations developed have been applied by making calculations for a Clark biplane and a Fairchild 22 monoplane" (p. 1).
From Summary: "This report describes tests conducted at the Langley Memorial Aeronautical Laboratory on an "air-inducting" exhaust manifold for aircraft engines. The exhaust gases from each cylinder port are discharged into the throat of an exhaust pipe which has a frontal bellmouth. Cooling air is drawn into the pipe, where it surrounds and mixes with the exhaust gases. Temperatures of the manifold shell and of the exhaust gases were obtained in flight for both a conventional manifold and the air-inducting manifold."
"An airplane in steady rectilinear flight was assumed to experience an initial disturbance in rolling or yawing velocity. The equations of motion were solved to see if it was possible to hasten recovery of a stable airplane or to secure recovery of an unstable airplane by the application of a single lateral control following an exponential law. The sample computations indicate that, for initial disturbances complex in character, it would be difficult to secure correlation with any type of exponential control" (p. 1).
Described here is an apparatus for studying the formation and combustion of fuel sprays under conditions closely simulating those in a high speed compression-ignition engine. The apparatus consists of a single-cylinder modified test engine, a fuel injection system so designed that a single charge of fuel can be injected into the combustion chamber, an electric driving motor, and a high-speed photographic apparatus. When the fuel is injected into the combustion chamber, motion pictures at the rate of 2000 per second are taken of the spray formation by means of spark discharges.
An investigation was made to determine the spinning characteristics of Clark Y monoplane wings with different plan forms. A rectangular wing and a wing tapered 5:2, both with rounded tips, were tested on the N.A.C.A. spinning balance in the 5-foot vertical wind tunnel. The aerodynamic characteristics of the models and a prediction of the angles of sideslip for steady spins are given. Also included is an estimate of the yawning moment that must be furnished by the parts of the airplane to balance the inertia couples and wing yawing moment for spinning equilibrium. The effects on the spin of changes in plan form and of variations of some of the important parameters are discussed and the results are compared with those for a rectangular wing with square tips. It is concluded that for a conventional monoplane using Clark Y wing the sideslip will be algebraically larger for the wing with the rounded tip than for the wing with the square tip and will be largest for the tapered wing. The effect of plan form on the spin will vary with the type of airplane; and the provision of a yawing-moment coefficient of -0.025 (i.e., opposing the spin) by the tail, fuselage, and interference effects will insure against the attainment of equilibrium on a steady spin for any of the plan forms tested and for any of the parameters used in the analysis.
An analysis of the factors governing the oscillation of an autogiro rotor blade in the plane of the rotor disk showed that the contribution of the air forces to the resultant motion was small and that the oscillation is essentially a direct effect of the rotor-blade flapping motion. A comparison of calculated oscillations with those measured in flight on three different rotors disclosed that the calculations gave satisfactory agreement with experiment. The calculated air forces on the rotor blade appear to be larger than the experimental ones, but this discrepancy can be attributed to the deficiencies in the strip analysis.
This note is one of a series covering an investigation of a family of related airfoils. It gives in preliminary form the results obtained from tests in the N.A.C.A. Variable-Density Wind Tunnel of two groups of six airfoils each. One group, the 43 series, has a maximum mean camber of 4 per cent of the chord at a position 0.3 of the chord from the leading edge; the other group, the 63 series, has a maximum mean camber of 6 per cent of the chord at the same position. The members within each group differ only in maximum thickness, the maximum thickness/chord ratios being:0.06, 0.09, 0.12, 0.15, 0.18, and 0.21. The results are analyzed with a view to indicating the variation of the aerodynamic characteristics with profile thickness for airfoils having a certain mean camber line.
"This note is one of a series covering an investigation of a number of related airfoils. It presents the results obtained from tests in N.A.C.A. Variable-Density Wind Tunnel of two groups of six airfoils each. One group, the 45 series, has a maximum mean camber of 4 per cent of the chord at a position 0.5 of the chord behind the leading edge, and the other group, the 65 series, has a maximum mean camber of 6 per cent of the chord at the same position. The members within each group differ only in maximum thickness, the maximum thickness/chord ratios being: 0.06, 0.09, 0.12, 0.15, 0.18, and 0.21. The results are analyzed with a view to indicating the variation of the aerodynamic characteristics with profile thickness for airfoils having a certain mean camber line form" (p. 1).
"Three airfoils, the N-85, the N-86, and the N-87, were tested at the request of the Bureau of Aeronautics, Navy Department, to determine the suitability of these sections for use as propeller-blade sections. Further tests of the NACA 0009-64 airfoil were also made to measure the aerodynamic effect of thickening the trailing edge in accordance with current propeller practice. The N-86 and the N-87 airfoils appear to be nearly equivalent aerodynamically and both are superior to the N-85 airfoil" (p. 1).
"This series of tests was undertaken to determine how much the reaction thrust of a jet could be increased by the use of thrust augmenters and thus to give some indication as to the feasibility of jet propulsion for airplanes. The tests were made during the first part of 1927 at the Langley Memorial Aeronautical Laboratory. A compressed air jet was used in connection with a series of annular guides surrounding the jet to act as thrust augmenters. The results show that, although it is possible to increase the thrust of a jet, the increase is not large enough to affect greatly the status of the problem of the application of jet propulsion to airplanes" (p. 1).
Note presenting an investigation in the 7- by 10-foot wind tunnel of an NACA 23021 airfoil with two arrangements of a 40-percent-chord slotted flap. The effect of slot shape, flap position, and flap deflection on the section aerodynamic characteristics was determined. Results regarding coefficients, precision, plain airfoils, and slotted-flap arrangements are provided.
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.
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