A report to the Weather Bureau, Washington DC, from the chairman of the Subcommittee on the Atmosphere in Relation to Aeronautics describing the activities accomplished and the proposal of work to be undertaken by the subcommittee.

From Summary: "Directly comparable drag measurements have been made of an airfoil with a conventional rectangular plan form and an airfoil with a sweptback plan form mounted on freely falling bodies. Both airfoils had NACA 65-009 sections and were identical in span, frontal area, and chord perpendicular to the leading edge. The sweptback plan form incorporated a sweepback angle of 45 degrees. The data obtained have been used to establish the relation between the airfoil drag coefficients and the free-stream Mach number over a range of Mach numbers from 0.90 to 1.27. The results of the measurements indicate that the drag of the sweptback plan form is less than 0.3 that of the rectangular plan form at a Mach number of 1.00 and is less than 0.4 that at a Mach number of 1.20."

From Summary :"A series of charts are presented by which the wing torsional stiffness required to meet a given standard of rolling effectiveness may be quickly determined. The charts may also be used to obtain quickly the aileron reversal speed and the variation of the loss in rolling effectiveness with airspeed. The charts apply to linearly tapered wings and elliptical wings of tubular-shell construction having various aspect ratios with aileron span and location of ailerons as variables. In the derivation of the charts, induced lift effects have been taken into account and the form of the wing-torsional-stiffness curve has been assumed."

"The study of combustion in a spark-ignition engine given in Technical Report no. 704 has been continued. The investigation was made with the NACA high-speed motion-picture camera and the NACA optical engine indicator. The camera operates at the rate of 40,000 photographs a second and makes possible the study of phenomena occurring in time intervals as short as 0.000025 second. Photographs are presented of combustion without knock and with both light and heavy knocks, the end zone of combustion being within the field of view" (p. 15).

Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of a large-chord wing model with a duct to house a simulated radiator suitable for a liquid-cooled engine. The duct was expanded to reduce the radiator losses, and the installation of the duct and radiator was made entirely within the wing to reduce form and interference drag. The tests were made using a two-dimensional-flow setup with a full-span duct and radiator. Section aerodynamic characteristics of the basic airfoil are given and also curves showing the characteristics of the various duct-radiator combinations.

A method is developed consistent with the assumptions of small perturbation theory which provides a means of determining the downwash behind a wing in supersonic flow for a known load distribution. The analysis is based upon the use of supersonic doublets which are distributed over the plan form and wake of the wing in a manner determined from the wing loading. The equivalence in subsonic and supersonic flow of the downwash at infinity corresponding to a given load distribution is proved.

"A method used by Tsien to derive similarity rules for hypersonic flows is utilized to derive Von Karman's similarity rules for transonic flows. At the lower limit of the transonic region of flow the theory yields a formula for the critical stream Mach numbers of a given family of symmetrical profiles. It is further shown that this formula can also be obtained by means of the Prandtl-Glauert small-perturbation method" (p. 83).

"Results are presented of an investigation that has been undertaken to develop theoretical methods of treating the motions of hydrofoil systems and to determine some of the important parameters. Variations of parameters include three distributions of area between the hydrofoils, two rates of change of downwash angle with angle of attack, three depths of immersion, two dihedral angles, two rates of change of lift with immersion, three longitudinal hydrofoil spacings, two radii of gyration in pitching, and various horizontal and vertical locations of the center of gravity. Graphs are presented to show locations of the center of gravity for stable motion, values of the stability roots, and motions following the sudden application of a vertical force or a pitching moment to the hydrofoil system for numerous sets of values of the parameters" (p. 1).

"Design charts are developed for 24S-T aluminum-alloy flat compression panels with longitudinal z-section stiffeners. These charts make possible the design of the lightest panels of this type for a wide range of design requirements. Examples of the use of the charts are given and it is pointed out on the basis of these examples that, over a wide range of design conditions, the maintenance of buckle-free surfaces does not conflict with the achievement of high structural efficiency" (p. 553).

Several methods of predicting the compressible-flow pressure loss across a baffled aircraft-engine cylinder were analytically related and were experimentally investigated on a typical air-cooled aircraft-engine cylinder. Tests with and without heat transfer covered a wide range of cooling-air flows and simulated altitudes from sea level to 40,000 feet. Both the analysis and the test results showed that the method based on the density determined by the static pressure and the stagnation temperature at the baffle exit gave results comparable with those obtained from methods derived by one-dimensional-flow theory. The method based on a characteristic Mach number, although related analytically to one-dimensional-flow theory, was found impractical in the present tests because of the difficulty encountered in defining the proper characteristic state of the cooling air. Accurate predictions of altitude pressure loss can apparently be made by these methods, provided that they are based on the results of sea-level tests with heat transfer.

"An investigation of the cooling of an 18-cylinder, twin-row, radial, air-cooled engine in a high-performance pursuit airplane has been conducted for variable engine and flight conditions at altitudes ranging from 5000 to 35,000 feet in order to provide a basis for predicting high-altitude cooling performance from sea-level or low altitude experimental results. The engine cooling data obtained were analyzed by the usual NACA cooling-correlation method wherein cylinder-head and cylinder-barrel temperatures are related to the pertinent engine and cooling-air variables. A theoretical analysis was made of the effect on engine cooling of the change of density of the cooling air across the engine (the compressibility effect), which becomes of increasing importance as altitude is increased" (p. 267).

From Summary: "A method is given for determining the horizontal tail loads in maneuvering flight. The method is based upon the assignment of a load-factor variation with time and the determination of a minimum time to reach peak load factor. The tail load is separated into various components. Examination of these components indicated that one of the components was so small that it could be neglected for most conventional airplanes; therefore, the number of aerodynamic parameters needed in this computation of tail loads was reduced to a minimum. In order to illustrate the method, as well as to show the effect of the main variables, a number of examples are given. Some discussion is given regarding the determination of maximum tail loads, maximum pitching accelerations, and maximum pitching velocities obtainable."

An experimental and analytical investigation of the flutter of sweptback cantilever wings is reported. The experiments employed groups of wings swept back by rotating and by shearing. The angle of sweep range from 0 degree to 60 degrees and Mach numbers extended to approximately 0.85. A theoretical analysis of the air forces on an oscillating swept wing of high length-chord ratio is developed, and the approximations inherent in the assumptions are discussed. Comparison with experiment indicates that the analysis developed in the present report is satisfactory for giving the main effects of sweep, at least for nearly uniform cantilever wings of high and moderate length-chord ratios.

Report presenting an investigation of the Reynolds number for transition on the outside of a hollow cylinder with heat transfer from the boundary layer to the wall at Mach number 6.9. At a given Mach number, it appears that the Reynolds number based on leading-edge thickness is an important parameter in comparisons of flat-plate transition data from various installations.

From Summary: "Various parts of aircraft propulsion engines that are in contact with hot gases often require cooling. Transpiration and film cooling, new methods that supposedly utilize cooling air more effectively than conventional convection cooling, have already been proposed. This report presents material necessary for a comparison of the cooling requirements of these three methods. Correlations that are regarded by the authors as the most reliable today are employed in evaluating each of the cooling processes."

Report presents the results of a full-scale investigation conducted in the NACA 20-foot tunnel to determine the pressure difference available for cooling with cowling flaps. The flaps were applied to an exit slot of smooth contour at 0 degree flap angle. Flap angles of 0 degree, 15 degrees, and 30 degrees were tested. Two propellers were used; propeller c which has conventional round blade shanks and propeller f which has airfoil sections extending closer to the hub. The pressure available for cooling is shown to be a direct function of the thrust disk-loading coefficient of the propeller.

Report presents the results of a study made of the effects of known end restraint on commercially available round and streamline tubing of chromium-molybdenum steel, duralumin, stainless steel, and heat-treated chromium-molybdenum steel; and a more accurate method than any previously available, but still a practical method, was developed for designing compression members in riveted or welded structures, particularly aircraft. Two hundred specimens were tested as short, medium-length, and long columns with freely supported ends or elastically restrained ends. Tensile and compressive tests were made on each piece of original tubing from which column specimens were cut.

"Fuel injected into a spherical bomb filled with air at a desired density and temperature could be ignited with a spark a few thousandths of a second after injection, an interval comparable with the ignition lag in fuel-injection engines. The effect of several variables on the extent and rate of combustion was investigated: time intervals between injection and ignition of fuel of 0.003 to 0.06 second and one of 5 minutes; initial air temperatures of 100 degrees C. to 250 degrees C.; initial air densities equivalent to 5, 10, and 15 absolute atmospheres pressure at 100 degrees C.; and air-fuel ratios of 5 to 25" (p. 107).

"Pressure measurements were made in flight on the right upper wing of an M-3 airplane. The effects of tip plan form, washout, and transverse camber were investigated with eight tip forms in unyawed conditions through the range of positive lift coefficients from zero lift to the stall. The conclusion is that the tip plan form does not influence the span distribution of the coefficients of normal force and moment. It is shown inferentially that temperature, humidity, and the aging of the wood and fabric wing structure used on the M-3 airplane have an appreciable influence on the load distribution" (p. 479).

This report presents the results of performance test conducted on a six-cylinder commercial fuel-injection pump that was adjusted to give uniform fuel distribution among the cylinders at a throttle setting of 0.00038 pound per injection and a pump speed of 750 revolutions per minute. The throttle setting and pump speed were then varied through the operating range to determine the uniformity of distribution and regularity of injection.

"This report describes the high-speed model towing basin of the National Advisory Committee for Aeronautics, usually referred to as the NACA Tank. The purpose of this piece of equipment is to enable the Committee to provide information and data regarding the performance of seaplanes on the water analogous to the information furnished concerning the performance of airplanes in the air" (p. 535).

This report presents the results of tests made in the 7-by 10-foot wind tunnel and in the 20-foot tunnel of the National Advisory Committee for Aeronautics to determine the drag of a number of airplane wheels, wheel fairings, and landing gears designed or selected for an airplane of 3,000 pounds gross weight. All tests were made on full-size models; those in the 7-by 10-foot tunnel were made at air speeds up to 80 miles per hour and those in the 20-foot tunnel were made at air speeds up to 100 miles per hour. Although most of the landing-gear tests were made in conjunction with a fuselage and at 0 degree pitch angle, some of the tests were made in conjunction with fuselage plus wings and a radial air-cooled engine and at pitch angles from -5 degrees to 6 degrees to obtain an indication of the general effect of these various items on landing-gear drag.

The data given in this report covers the explosive limits of hydrocarbon fuels. Incidental to the purpose of the investigation here reported, the explosive limits will be found to be expressed for the condition of constant pressure, in the fundamental terms of concentrations (partial pressures) of fuel and oxygen.

This report outlines the development of a general theory for the calculation of the effect of the boundaries of the air stream on the flow past an airfoil. Analytical treatments are given for tunnels with horizontal boundaries only, with vertical boundaries only, and with a bottom boundary only. Formulas are developed for the tunnel wall interference in each case for an airfoil located at the center of the tunnel. The correction is given as a function of the width to height ratio of the tunnel.