Report presenting large-scale tests of a semispan model of a 60 degree sweptback wing-fuselage combination. The model was equipped with 25-percent-chord, 36-percent-span, double-slotted flaps, 25-percent-chord, 43-percent-span ailerons, and 17.5-percent-chord leading-edge slats. Flow separation limited the linear range of the pitching moment of the base wing to a lift coefficient of about 0.35.

From Summary: "By application of a hyperbolic approximation to the form of the bow waves caused by blunt leading edges on an infinite cascade of supersonic blades, the approximate losses in relative total pressure due to the external bow-wave system arising from blunt edges and subsonic axial entrance velocities were computed. The losses increase linearly with leading-edge radius for any given relative Mach number. For a relative Mach number of 1.60, leading-edge radii may be approximately 1.5 percent of the normal blade gap with a 1-percent loss of relative total pressure."

Report presenting a method for calculating the effect of the fuselage on the spanwise lift distribution on a wing by applying a conformal-mapping procedure to the simplified lifting-surface theory. The calculations are also compared with experimental data and are found to generally agree.

An investigation was made of the change in the pressure limits of flame propagation with tube diameter for various isooctane-oxygen-nitrogen mixtures. Pressure limits were measured in cylindrical glass tubes of four different inside diameters at six different oxygen-nitrogen ratios. Under the experimental conditions, flame propagation was found to be impossible in isooctane-oxygen mixtures with oxygen concentrations less than 11 to 12 percent. Critical tube diameters for flame propagation were calculated and the effect of pressure was determined and compared with the effect of pressure on quenching distance. Critical diameters were related to flame speeds for various isooctane-oxygen-nitrogen mixtures.

The experimental performance of turbine nozzle blades designed for a constant discharge angle was investigated at discharge hub Mach numbers of 1.18, 1.31, and 1.41. Flow characteristics are presented in terms of energy losses, angle gradients, and secondary flow effects. Blade efficiency decreased from 0.983 to 0.978 with increasing Mach number in the range investigated while angle variations in the loss regions became very large, indicating poorer blade performance than efficiency implies.

Report presenting experimental damping-in-roll derivatives obtained for a series of 33 swept and tapered wings. The wing plan forms were selected so that a range of leading-edge positions ahead of and behind the Mach cone obtained at three Mach numbers.

Report presenting a method for analyzing each of the elements, boron, carbon, and hydrogen from the quantitative analysis of a single sample of an organoboron compound. The method is helpful for the analysis of volatile hydrocarbons and polynuclear aromatic compounds.

From Summary: "An experimental wind-tunnel investigation was conducted at Mach 1.9 to determine the pressures acting on the base of a multijet missile using unheated air and carbon dioxide as jet fluids. The variation of base pressure with jet static-pressure ratio was compared with results estimated for an axisymmetric single-jet model and some correlation was observed."

Memorandum presenting an investigation of the effective thermal conductivities of powders, which included uranium oxide UO(sub 2) powder with a density of 59.5 percent of the density of the solid uranium oxide at a range of temperatures. Helium, argon, and nitrogen gases and mixtures of helium-argon and xenon-krypton gases were introduced into the powder void, and the experimental effective thermal conductivities of the gas-powder mixtures were determined.

Report presenting an experimental investigation to determine the internal performance characteristics of a series of circular, open-nose auxiliary air inlets immersed into the supersonic, turbulent boundary layer generated by the 8- by 6-foot supersonic wind tunnel. Inlet critical total-pressure and mass-flow ratios were reasonably predicted by theory at all conditions investigated. Results regarding boundary-layer flow, inlet performance, and diffuser-exit pressure fluctuations are provided.

Report presenting an investigation to study the effects of several heat treatments on the operating life of turbine blades in a J33-33 turbojet engine operated without an afterburner. As-forged blades, blades solution treated at temperatures high enough to produce germinated grains, blades given a double-aging treatment, and blades overaged by overtemperature heat treatments are evaluated. Results regarding the engine operating results and blade performance, metallurgical studies of as-heat-treated specimens, and metallurgical studies of failed blades are provided.

"A method is described by which the total temperature of the gases at the combustion-chamber outlet of a ram-jet engine may be determined from the loss in total pressure measured across the combustion chamber. A working chart is presented by means of which the ratio of the total temperature of the gases at the combustion-chamber outlet to the total temperature of the gases at the combustion-chamber inlet may be determined from the measured loss of total pressure across the combustion chamber and the known values of air flow, total pressure, and total temperature at the combustion-chamber inlet. Values of total-temperature ratio across the combustion chamber of a 20-inch ram jet were obtained in the Cleveland altitude wind tunnel over a range of pressure altitudes from 6000 to 15,000 feet" (p. 1).

"Static longitudinal and lateral stability characteristics of a swept-wing fighter-bomber airplane model with a top inlet have been determined experimentally. In addition, vertical-tail effectiveness, spoiler effectiveness, effects of several store configurations, and effects of mass flow were investigated. Tests were made at Mach numbers of 1.6, 1.8, 2.0, 2.2, and 2.35 and Reynolds numbers, based upon mean aerodynamic chord, of between 2.0x10(exp 6) and 2.5x10(exp 6)" (p. 1).

Report presenting flight tests with an airplane equipped to vary the longitudinal-control-system dynamic characteristics over a wide range. Control is obtained through a stabilizer force-command position-feedback longitudinal-control system. Results at a Mach number of 0.80 at 35,000 feet altitude, Mach number of 0.35 at 5,000 feet altitude, and normal-acceleration response are provided.

Report presenting values of zero angle-of-attack performance of a double-cone inlet with and without a shroud through a range of mass-flow ratios and Mach numbers. Maximum propulsive thrust was computed to be higher with the shrouded inlet at all supersonic Mach numbers tested. Results regarding internal diffuser performance, drag, and engine performance are provided.

"A summary of the locations, extents, and intensities of turbulent areas encountered during a flight investigation of the jet stream is presented. The data were obtained from NACA VGH time-history records" (p. 1).