Report presenting an analytical investigation to determine the effect of compressor interstage air bleed with the use of constant-area bleed ports on the acceleration characteristics of a typical high-pressure-ratio single-spool turbojet engine. Constant-area interstage bleed, properly located, gave smaller acceleration times than variable-area compressor exit bleed. Results regarding acceleration with interstage bleed, acceleration using constant-area interstage bleed in combination with compressor outlet bleed, variable-area interstage bleed, and comparison of acceleration modes are provided.

From Introduction: "The purpose of the present paper was to determine whether the characteristics about a skewed axis could be predicted from data about the normal angle-of-attack axis, and whether such a configuration offered any aerodynamic advantages over the conventional hinge location normal to the pane of symmetry."

"Temperature surveys were made at the combustion-chamber outlets of a British Rolls-Royce Nene II engine. The highest mean nozzle-vane and mean gas temperatures were found to occur at a radius approximately 75% of the nozzle-vane length from the inner ring of the nozzle-vane assembly. Variations in engine speed, jet-nozzle area, simulated altitude, and simulated flight speed altered the temperature level but did not materially affect the pattern of radial temperature distribution" (p. 1).

Report presenting an investigation to determine the effect of conical camber on the drag due to lift and on the static longitudinal, lateral, and directional stability characteristics of an aspect-ratio-3, 45 degree sweptback wing in combination with a streamline body. The addition of conical camber to the basic wing reduced the drag due to lift at moderate and large lift coefficients and increased the maximum lift-drag ratios.

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.

Report presenting an investigation at conditions of zero ram and seal-level pressure to determine the effect of inlet temperature and humidity on turbojet engine performance with injection of water and water-alcohol mixtures at the compressor inlet.

Report presenting a wind-tunnel investigation to determine the effects of leading-edge radius and maximum thickness-chord ratio on the variation with Mach number of the aerodynamic characteristics of several thin symmetrical NACA 4-digit-series airfoil sections. The variations with Mach number of the lift, drag, and pitching moment for a 4-percent-chord-thick airfoil section are not significantly affected by a change of leading-edge radius from 0.18 to 0.53 percent of the chord. Results regarding the leading-edge radius effects and maximum thickness-chord ratio are provided.

Report presenting an experimental investigation of thrust augmentation by bleedoff using a 4000-pound-thrust turbojet engine at zero flight-speed, sea-level conditions. Secondary combustion air was bled off from combustion chambers and discharged from an exhaust nozzle and water was then injected into the combustion chambers to replace the bleed-off air. Results regarding the thrust characteristics, engine power, and augmented thrust ratio are provided.

Report presenting a flight investigation of a rocket-launched solid-fuel ram-jet engine designed to bypass cooling air around the fuel charge. Having an internally cooled combustor averted combustor burn-out during the flight test. Results regarding acceleration, altitude, air specific impulse, and fuel specific impulse are provided.

Report presenting a flight investigation of a rocket-launched solid-fuel ram-jet engine designed to bypass cooling air around the fuel charge. The internally cooled combustor averted combustor burn-out during the flight test. Results regarding the net thrust coefficient, gross thrust coefficient, time history of the air specific impulse, and total fuel load are provided.

Memorandum presenting an investigation of a 14-inch axial-flow-compressor rotor with circular-arc blades operating up to a rotor-inlet relative Mach number of 1.22 in order to determine the blade-element characteristics. Results regarding overall performance, blade-element performance, performance of rotor with guide vanes, performance of rotor alone, and comparison of rotor investigation are provided.

Report presenting the performance of the liquid fluorine-liquid ammonia propellant combination investigated in 100-pound-thrust, water-cooled engines operating at a chamber pressure of 300 pounds per square inch absolute. Several impinging-jet injectors were evaluated in chambers of characteristic length equal to 50 inches. Results regarding specific impulse, characteristic velocity, thrust coefficient, heat rejection, and operational notes are provided.

"A preliminary evaluation of the spin and recovery characteristics of the XF3D-1 airplane has been made, based primarily on the results of the free-spinning tunnel tests of a model which closely simulated the XF3D-1 in tail design, tail length, and mass loading. Estimates have been made of the rudder-pedal force that may be encountered in effecting recovery from a spin and of the spin recovery parachute requirements of the airplane for demonstration spins. The method of bail-out which should be used if it becomes necessary for the crew to abandon the airplane during a spin is indicated" (p. 1).

Memorandum presenting a preliminary investigation to determine the effects of alpha-particle bombardment on the creep rate of aluminum wire at 400 degrees Fahrenheit. The alpha radiation from an 85-millicurie polonium source appeared to decrease slightly the creep rate of the aluminum.

Memorandum presenting an instrument for recording and indicating the frequency and intensity of aircraft icing conditions to obtain statistical icing data over worldwide air routes during routine airline operations. Data of this type is needed to provide pertinent meteorological information necessary for the optimum design of ice-protection systems. The features that make this particular instrument desirable include the simplicity of operation, freedom from maintenance and operating problems, automatic operation upon encountering an icing condition, visual indications of icing intensity available to the pilot, total weight of 18 pounds, and continuous recorded data.

"An instrument for recording and indicating the frequency and intensity of aircraft icing conditions encountered in flight has been developed by the NACA Lewis Laboratory to obtain statistical icing data over world-wide air routes during routine airline operations. The operation of the instrument is based on the creation of a differential pressure between an ice-free total-pressure system and a total-pressure system in which small total-pressure holes vented to static pressure are allowed to plug with ice accretion. The simplicity of this operating principle permits automatic operation, and provides relative freedom from maintenance and operating problems" (p. 1).

Memorandum presenting certain considerations and techniques pertinent to the application of the supersonic area rule to the design of airplane fuselages. Some of the factors considered are an extension of the rule to account for the interference effects of the wing and tail on the general flow field for asymmetrical configurations, the determination of fuselage area developments which result in approximately the minimum wave drag, and the influence of wing parameters and design Mach number on the effectiveness of fuselage shaping.

Report presenting some considerations and techniques pertinent to the application of the supersonic area rule to the design of airplane fuselages. Some of the factors to consider are an extension of the rule to account for the interference effects of the wing and tail on general flow field, the determination of fuselage area developments which result in minimum wave drag, and the influence of wing parameters and Mach number on the effectiveness of fuselage shaping.

"Theoretical values of performance parameters for liquid ammonia and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity" (p. 1).

"Theoretical values of performance parameters for liquid ammonia and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity" (p. 1).

Report presenting aerodynamic-heating data obtained from Mach numbers 1.2 to 4.2 from a free-flight test of a model finless NACA RM-10. Skin temperatures were measured at six stations along the body. Results regarding experimental recovery factors and Stanton number are provided.