National Advisory Committee for Aeronautics (NACA) - 78 Matching Results

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Abstracts pertaining to seaplanes

Description: Report discussing about 400 references pertaining to the hydrodynamic design of seaplanes have been compiled, and the information is presented in the form of abstracts classified under six main headings.
Date: July 24, 1947
Creator: Bidwell, Jerold M & King, Douglas A
Partner: UNT Libraries Government Documents Department

Analytical Treatment of Normal Condensation Shock

Description: The condensation of water vapor in an air consequences: acquisition of heat (liberated heat vaporization; loss of mass on the part of the flowing gas (water vapor is converted to liquid); change in the specific gas constants and of the ratio k of the specific heats (caused by change of gas composition). A discontinuous change of state is therefore connected with the condensation; schlieren photographs of supersonic flows in two-dimensional Laval nozzles show two intersecting oblique shock fronts that in the case of high humidities may merge near the point of intersection into one normal shock front.
Date: July 1, 1947
Creator: Heybey
Partner: UNT Libraries Government Documents Department

Calculation of the Pressure Distribution on Bodies of Revolution in the Subsonic Flow of a Gas, Part 1, Axially Symmetrical Flow

Description: The present report concerns a method of computing the velocity and pressure distributions on bodies of revolution in axially symmetrical flow in the subsonic range. The differential equation for the velocity potential Phi of a compressible fluid motion is linearized tn the conventional manner, and then put in the form Delta(Phi) = 0 by affine transformation. The quantity Phi represents the velocity potential of a fictitious incompressible flow, for which a constant superposition of sources by sections is secured by a method patterned after von Karman which must comply with the boundary condition delta(phi)/delta(n) = 0 at the originally specified contour. This requirement yields for the "pseudo-stream function" psi a differential equation which must be fulfilled for as many points on the contour as source lengths are assumed. In this manner, the problem of defining the still unknown source intensities is reduced to the solution of an inhomogeneous equation system. The pressure distribution is then determined with the aid of Bernoulli's equation and adiabatic equation of state. Lastly, the pressure distributions in compressible and incompressible medium are compared on a model problem.
Date: July 1, 1947
Creator: Bilharz, Herbert & Hoelder, Ernst
Partner: UNT Libraries Government Documents Department

Considerations of the Total Drag of Supersonic Airfoil Sections

Description: The results of calculations of the viscous and pressure drags of some two-dimensional supersonic airfoils at zero lift are presented. The results indicate that inclusion of viscous drag alters many previous results regarding the desirability of certain airfoil shapes for securing low drags at supersonic speeds. At certain Reynolds and Mach numbers, for instance, a circular-arc airfoil may theoretically have less drag than the previously advocated symmetrical wedge-shape profile; although under different conditions, the circular-arc airfoil may have a higher drag.
Date: July 1947
Creator: Ivey, H. Reese & Klunker, E. Bernard
Partner: UNT Libraries Government Documents Department

Distribution of wave drag and lift in the vicinity of wing tips at supersonic speeds

Description: The point-source-distribution method of calculating the aerodynamic coefficients of thin wings at supersonic speeds was extended to include the effect of the region between the wing boundary and the foremost Mach wave from the wing leading edge. The effect of this region on the surface velocity potential has been determined by an equivalent function, which is evaluated over a portion of the wing surface. In this manner, the effect of angles of attack and yaw as well as the asymmetry of top and bottom wing surfaces may be calculated. As examples of the method, the pressure distribution on a thin plate wing of rectangular plan form as well as the lift and the drag coefficients as a function of Mach number, angle of attack, and aspect ratio are calculated. The equations for the surface velocity potential of several other plan forms are also included. (author).
Date: July 1, 1947
Creator: Evvard, John C
Partner: UNT Libraries Government Documents Department

Drag Corrections in High-Speed Wind Tunnels

Description: In the vicinity of a body in a wind tunnel the displacement effect of the wake, due to the finite dimensions of the stream, produces a pressure gradient which evokes a change of drag. In incompressible flow this change of drag is so small, in general, that one does not have to take it into account in wind-tunnel measurements; however, in compressible flow it beoomes considerably larger, so that a correction factor is necessary for measured values. Correction factors for a closed tunnel and an open jet with circular cross sections are calculated and compared with the drag - corrections already bown for high-speed tunnnels.
Date: July 1, 1947
Creator: Ludwieg, H.
Partner: UNT Libraries Government Documents Department

Drag Reduction by Suction of the Boundary Layer Separated Behind Shock Wave Formation at High Mach Numbers

Description: With an approach of the velocity of flight of a ship to the velocity of sound, there occurs a considerable increase of the drag. The reason for this must be found in the boundary layer separation caused by formation of shock waves. It will be endeavored to reduce the drag increase by suction of the boundary layer. Experimental results showed that drag increase may be considerably reduced by this method. It was, also, observed that, by suction, the position of shock waves can be altered to a considerable extent.
Date: July 1, 1947
Creator: Regenscheit, B.
Partner: UNT Libraries Government Documents Department

Effect of fuel on performance of a single combustor of an I-16 turbojet engine at simulated altitude conditions

Description: As part of a study of the effects of fuel composition on the combustor performance of a turbojet engine, an investigation was made in a single I-16 combustor with the standard I-16 injection nozzle, supplied by the engine manufacturer, at simulated altitude conditions. The 10 fuels investigated included hydrocarbons of the paraffin olefin, naphthene, and aromatic classes having a boiling range from 113 degrees to 655 degrees F. They were hot-acid octane, diisobutylene, methylcyclohexane, benzene, xylene, 62-octane gasoline, kerosene, solvent 2, and Diesel fuel oil. The fuels were tested at combustor conditions simulating I-16 turbojet operation at an altitude of 45,000 feet and at a rotor speed of 12,200 rpm. At these conditions the combustor-inlet air temperature, static pressure, and velocity were 60 degrees F., 12.3 inches of mercury absolute, and 112 feet per second respectively, and were held approximately constant for the investigation. The reproducibility of the data is shown by check runs taken each day during the investigation. The combustion in the exhaust elbow was visually observed for each fuel investigated.
Date: July 3, 1947
Creator: Zettle, Eugene V; Bolz, Ray E & Dittrich, R T
Partner: UNT Libraries Government Documents Department

Effects of sweepback on boundary layer and separation

Description: Following the law of stress adopted in the Navier-Stokes equations, the configuration of the viscous flow in planes at right angles to the axis of an infinite cylinder is found to be independent of the axial motion of the cylinder. In the limiting case of a yawed or swept wing of very high aspect ratio, certain boundary-layer and separation phenomena are thus determined independently by the crosswise component of velocity.
Date: July 1, 1947
Creator: Jones, Robert T
Partner: UNT Libraries Government Documents Department