This report presents the results of wind tunnel tests of two airship models conducted to determine the drag coefficients at zero pitch, and the effect of fins and cars and of flat and streamlined protuberances located at various positions along the hull. During the investigation the stern of one model was rounded off to produce a blunter shape. The extreme range of the Reynolds number based on the over-all length of the models was from 1,300,000 to 33,000,000. At large values of the Reynolds number the streamlined protuberance affected the drag very little, and the additional drag caused by the flat protuberance was less than the calculated drag by the protuberance alone. The fins and cars together increased the bare-hull drag about 20 per cent.

"An investigation of the nature of the flow field behind a rectangular wing of circular arc cross section has been conducted in the Langley 9-inch supersonic tunnel. Pitot- and static-pressure surveys covering a region of flow behind the wing have been made together with detailed pitot surveys throughout the region of the wake. In addition, the flow direction has been measured by means of a weathercocking vane. Theoretical calculations have been made to obtain the variation of both downwash and sidewash with angle of attack by using the superposition method of Lagerstrom, Graham, and Grosslight" (p. 1).

This report presents the results of flight measurements of longitudinal stability and control characteristics made on a swept-wing jet aircraft to determine the origin of the pitch-up encountered in maneuvering flight at transonic speeds. For this purpose measurements were made of elevator angle, tail angle of attack, and wing-fuselage pitching moments (obtained from measurements of the balancing tail loads).

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.

"A method is described for measuring the dynamic stability derivatives of a model airplane in a wind tunnel. The characteristic features of this system are that single-degree-of-freedom oscillations were used to obtain combinations of rolling, yawing and pitching motions; that the oscillations were excited and controlled by velocity feedback which permitted operation under conditions unfavorable for more conventional types of oscillatory testing; and that data processing was greatly simplified by using analog computer elements in the strain-gage circuitry. A small number of experimental data are included to illustrate the general scope of results obtainable with this system" (p. 1).

In connection with the general problem of providing air flow to an aircraft power plant located within a fuselage, an investigation was conducted in the Langley 8-foot high-speed tunnel to determine the effect on external drag and pressure distribution of air inlet openings located at the nose of a streamline body. Air outlet openings located at the tail and at the 21-percent and 63-percent stations of the body were also investigated. Boundary layer transition measurements were made and correlated with the force and the pressure data. Individual openings were investigated with the aid of a blower and then practicable combinations of inlet and outlet openings were tested. Various modifications to the internal duct shape near the inlet opening and the aerodynamic effects of a simulated gun in the duct were also studied. The results of the tests suggested that outlet openings should be designed so that the static pressure of the internal flow at the outlet would be the same as the static pressure of the external flow in the vicinity of the opening.

"A single-stage axial fan was built and tested in the shop of the propeller-research tunnel of the NACA. The fan comprised a simple 24-blade rotor having a diameter of 21 inches and a solidity of 0.86 and a set of 37 contravanes having a solidity of 1.33. The rotor was driven by a 25-horsepower motor capable of rotating at a speed of 3600 r.p.m. The fan was tested for volume, pressure, and efficiency over a range of delivery pressures and volumes for a wide range of contravane and blade-angle settings" (p. 49).

"Evaluation of the rotating multicylinder method for the measurement of droplet-size distribution, volume-median droplet size, and liquid-water content in clouds showed that small uncertainties in the basic data eliminate the distinction between different cloud droplet-size distributions and are a source of large errors in the determination of the droplet size. Calculations of the trajectories of cloud droplets in incompressible and compressible flow fields around a cylinder were performed on a mechanical analog constructed for the study of the trajectories of droplets around aerodynamic bodies" (p. 1).

From Summary: "A simplified lifting-surface theory is applied to the problem of evaluating span loading due to flap deflection for arbitrary wing plan forms. With the resulting procedure, the effects of flap deflection on the span loading and associated aerodynamic characteristics can be easily computed for any wing which is symmetrical about the root chord and which has a straight quarter-chord line over the wing semispan. The effects of compressibility and spanwise variation of section lift-curve slope are taken into account by the procedure. The method presented can also be used to calculate the downwash in the vertical center of the wake of a wing which has arbitrary spanwise loading."

This report contains the derivation of a series of simple approximate equations for density ratios and for the pressure ratio in the standard atmosphere. The accuracy of the various equations is discussed and the limits of applications are given. Several of these equations are in excellent agreement with the standard values.

Static thrust data from more than 100 airplane propeller tests are collected from various sources and combined in working charts, from which the static thrust coefficient may be readily determined. The available data cover practically all types of propellers and are in good agreement.

From Summary: "It is demonstrated with the aid of similitude arguments that a model launched from a hypervelocity gun upstream through a special supersonic nozzle should experience aerodynamic heating and resulting thermal stresses like those encountered by a long-range ballistic missile entering the earth's atmosphere. This demonstration hinges on the requirements that model and missile be geometrically similar and made of the same material, and that they have the same flight speed and Reynolds number (based on conditions just outside the boundary layer) at corresponding points in their trajectories. The hypervelocity gun provides the model with the required initial speed, while the nozzle scales the atmosphere, in terms of density variation, to provide the model with speeds and Reynolds numbers over its entire trajectory."

The correlation functions and power spectra of the rolling and yawing moments on an airplane wing due to the three components of continuous random turbulence are calculated. The rolling moments to the longitudinal (horizontal) and normal (vertical) components depend on the spanwise distributions of instantaneous gust intensity, which are taken into account by using the inherent properties of symmetry of isotropic turbulence. The results consist of expressions for correlation functions or spectra of the rolling moment in terms of the point correlation functions of the two components of turbulence.

"Following a brief history of the NACA investigation of jet propulsion, a discussion is given of the general investigation and analysis leading to the construction of the jet-propulsion ground-test mock-up. The results of burning experiments and of test measurements designed to allow quantitative flight performance predictions of the system are presented and correlated with calculations. These calculations are then used to determine the performance of the system on the ground and in the air at various speeds and altitudes under various burning conditions. The application of the system to an experimental airplane is described and some performance predictions for this airplane are made" (p. 1).

The properties of conical flow without axial symmetry are analyzed. The flow around cones of circular cross section at small angles of attack is determined by correctly considering the effect of the entropy gradients in the flow.

"Simultaneous measurements were made of the speed of flame and the rise in pressure during explosions of mixtures of carbon monoxide, normal heptane, iso-octane, and benzene in a 10-inch spherical bomb with central ignition. From these records, fundamental properties of the explosive mixtures, which are independent of the apparatus, were computed. The transformation velocity, or speed at which flame advances into and transforms the explosive mixture, increases with both the temperature and the pressure of the unburned gas" (p. 39).

Report presents the results of an investigation conducted to study the effects of the inert gases, argon and helium, upon flame speed and expansion ratio in exploding mixtures of carbon monoxide, oxygen and water. For the particular gas mixtures investigated the results show that: (1) With the possible exception of helium in small amounts the addition of inert gas always produces decreased flame speed and expansion ratio; (2) like volumes of argon and helium have very different effects upon flame speed but practically the same effect upon expansion ratio; and (3) the difference in the effect of these two gases upon speed is independent of the ratio of carbon monoxide to oxygen. A discussion of some possible modes by which inert gases may produce the observed effects is included.

From Summary: "The differential equation of Chaplygin's jet problem is utilized to give a systematic development of particular solutions of the hodograph flow equations, which extends the treatment of Chaplygin into the supersonic range and completes the set of particular solutions. The particular solutions serve to place on a reasonable basis the use of velocity correction formulas for the comparison of incompressible and compressible flows. It is shown that the geometric-mean type of velocity correction formula introduced in part I has significance as an over-all type of approximation in the subsonic range."

From Summary: "The differential equation of Chaplygin's jet problem is utilized to give a systematic development of particular solutions of the hodograph flow equations, which extends the treatment of Chaplygin into the supersonic range and completes the set of particular solutions. The particular solutions serve to place on a reasonable basis the use of velocity correction formulas for the comparison of incompressible and compressible flows. It is shown that the geometric-mean type of velocity correction formula introduced in part I has significance as an over-all type of approximation in the subsonic range."

"This report presents the results of an investigation to determine the coefficients of discharge of nozzles with small, round orifices of the size used with high-speed compression-ignition engines. The injection pressures and chamber back pressures employed were comparable to those existing in compression-ignition engines during injection. The construction of the nozzles was varied to determine the effect of the nozzle design on the coefficient. Tests were also made with nozzles assembled in an automatic injection valve, both with a plain and with a helically grooved stem" (p. 193).

This report presents the results of an investigation conducted to determine the interrelation of the constituents of the exhaust gases of internal-combustion engines and the effect of engine performance on these relations. Six single-cylinder, liquid-cooled tests engines and one 9-cylinder radial air-cooled engine were tested. Various types of combustion chambers were used and the engines were operated at compression ratios from 5.1 to 7.0 using spark ignition and from 13.5 to 15.6 using compression ignition. The investigation covered a range of engine speeds from 1,500 to 2,100 r.p.m.

In order to provide engineers interested in rotating-wing aircraft, but with no specialized training in stability theory, some understanding of the factors that influence the flying qualities of the helicopter, an explanation is made of both the static stability and the stick-fixed oscillation in hovering and forward flight in terms of fundamental physical quantities. Three significant stability factors -- static stability with angle of attack, static stability with speed, and damping due to a pitching or rolling velocity -- are explained in detail.

"Existing conical-flow solutions have been used to calculate the hinge-moments and effectiveness parameters of trailing-edge controls having leading and trailing edges swept ahead of the Mach lines and having streamwise root and tip chords. Equations and detailed charts are presented for the rapid estimation of these parameters. Also included is an approximate method by which these parameters may be corrected for airfoil-section thickness" (p. 937).

"From consideration of available information on boundary-layer behavior, a relation among profile thickness, maximum surface velocity, Reynolds number, velocity diagram, and solidity is established for a cascade of airfoils immersed in a two-dimensional incompressible fluid flow. Several cascades are computed to show the effect of various cascade design parameters on minimum required cascade solidity. Comparisons with experimentally determined blade performance show that the derived blade loadings are equal or higher for moderate flow deceleration and somewhat lower for large deceleration. Blades with completely laminar flow appear practical for impulse or reaction blading" (p. 117).