Memorandum presenting an investigation of variable-area auxiliary inlets intended to provide for the variation of engine air requirements with speed and altitude. Three auxiliary scoop inlets were tested in combination with a nose inlet at a range of Mach numbers tested at zero angle of attack.

Report presenting free-flight and free-jet testing of a twin side-inlet configuration, including a pilot's canopy and a wheel-well fairing, between a range of Mach numbers from 0.8 to 1.55. Using side inlets provides two important advantages: the inlet can be located closer to the engine and it frees up the nose for other uses. Results regarding the effects of oscillating flow, external-drag coefficient, and total-pressure recovery are provided.

Report presenting an experimental investigation of pressure disturbances caused by the reflection of two-dimensional shock waves from a slotted boundary at Mach numbers between 1.0 and 1.5. Simple-wedge airfoils with three different apex angles were used to generate shock waves and the static pressures ere measured at various points along the surface of each airfoil.

Report presenting testing of a multiweb wing structure, representing an airplane or missile wing, under simulated supersonic flight conditions to determine the transient temperature distribution. The aerodynamic loads caused the model to fail near the end of the test. The failure is analyzed and it is determined that the model failed as a result of the combined action of aerodynamic heating and loading.

Report presenting a connected-pipe investigation to evaluate a can-type combustor configuration and to develop the configuration to give a wide operable fuel-air ratio range of high combustion efficiency. Results regarding the wind-tunnel tests and connected-pipe tests are provided.

Report presenting theoretical values of rocket performance parameters for two mixtures of liquid ammonia and hydrazine with flourine as oxidant for a variety of parameters. Parameters included are 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.

Report presenting theoretical values of rocket performance parameters for two mixtures of liquid ammonia and hydrazine as fuels with liquid fluorine as oxidant. 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.

Report presenting an investigation to determine the effects of geometric dihedral on the aerodynamic characteristics of wing-fuselage combinations with wings of aspect ratio 4, taper ratio 0.6, and angles of sweep of 3.6 and 45 degrees at the quarter-chord line. Little analysis is provided, but the results indicate that the effect of geometric dihedral on the effective-dihedral parameter is slightly larger than predicted at angles of attack of about 6 degrees.

From a theory developed on a quasi-one-dimensional-flow basis, it is found that the stability of the ram jet is dependent upon the instantaneous values of mass flow and total pressure recovery of the supersonic diffuser and immediate neighboring subsonic diffuser. Conditions for stable and unstable flow are presented. The theory developed in the report is in agreement with the experimental data of NACA-TN-3506 and NACA-RM-L50K30. A simple theory for predicting the approximate amplitude of small pressure pulsation in terms of mass-flow decrement from minimum-stable mass flow is developed and found to agree with experiments.

"An investigation has been made in the NACA Lewis icing research tunnel to determine the aerodynamic and icing characteristics of a full-scale induction-system air-scoop assembly incorporating a flush alternate inlet. The flush inlet was located immediately downstream of the offset ram inlet and included a 180 deg reversal and a 90 deg elbow in the ducting between inlet and carburetor top deck. The model also had a preheat-air inlet" (p. 1).

Memorandum presenting an investigation in the icing research tunnel to determine the aerodynamic and icing characteristics of a full-scale induction-system air-scoop assembly incorporating a flush alternate inlet. Results regarding the ram-inlet tests, alternate-inlet tests, and preheat-air tests are provided.

An experimental investigation has been conducted in the Langley stability tunnel at low speed to determine the pitching stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient, control deflections, and propeller blade angle were investigated. The tests were made through an angle-of-attack range from about -4deg to 29deg, and the thrust coefficient range was from 0 to 0.7. In order to expedite distribution of these data, no analysis of the data has been prepared for this paper.

Report presenting results of a systematic study in the 8- by 6-foot supersonic wind tunnel to determine the effect that area distribution in the subsonic diffuser might have on stability and overall performance of conical-type supersonic diffusers. Drag, pressure recovery, and mass flow are presented for five diffusers designed for maximum mass flow at Mach number 2.0.

Report presenting a series of turbine wheel designs made for a turbojet engine with an axial-flow compressor. The designs were made to determine the influence of air-cooling on overall engine design, weight, critical-material content, and complexity and to assist with construction of research model full-scale turbines.

"Eighty pounds of butylsilanes required for test purposes was synthesized by reducing butyltrichlorosilane with lithium aluminum hydride in dioxane solution. This was safely and successfully accomplished by operating in steel equipment under an atmosphere of oil-pumped nitrogen to minimize the hazards involved in conducting this operation on a large scale. The apparatus and procedure used to carry out the necessary manipulations are described in this report" (p. 1).

Memorandum presenting a summary of the methods for estimation of water landing loads on hydro-ski-equipped aircraft in order to provide a unified picture and bibliography of NACA work on the subject. Two approaches to the problem of calculating hydro-ski landing loads are covered. Results regarding the impact load theory for rigidly mounted hydro-skis, effect of chine immersion on landing loads, impact load theory for shock-mounted hydro-skis, and estimation of the water pressure distribution on impacting hydro-skis are provided.

"An investigation was conducted in the Lewis altitude wind tunnel to evaluate the performance and operational characteristics of the J40-WE-8 afterburner. A brief program of minor modifications to the flame holder, diffuser, and fuel system was undertaken to improve at a burner-inlet pressure level of 620 pounds per square foot. At this pressure level, modifications to the fuel system resulted in an increase in maximum net thrust from 1500 to 1600 pounds and a reduction in specific fuel consumption in the stoichiometric region from 3.70 to 3.15 pounds of fuel per hour per pound of net thrust" (p. 1).

Report presenting burning rates for single isooctane drops suspended in various quiescent oxygen-nitrogen atmospheres at room temperature and pressure. The burning rates were compared with those predicted by a previously developed theory based on a heat- and mass-transfer mechanism and with values predicted by a modification to the theory.

Memorandum presenting an investigation of the durabilities of several protective coatings applied to air-cooled gas turbine rotor blades of nonstrategic steels in modified turbojet engines. Four types of coatings (ceramic, nickel, Nicrobraz, and aluminized) were applied to a total of 20 blades.

A program was undertaken to determine the J73 turbojet engine compressor stall and surge characteristics and combustor blow-out limits encountered during transient engine operation. Data were obtained in the form of oscillograph traces showing the time history of several engine performance parameters with changes in engine fuel flow. The data presented in this report are for step changes in fuel flow at an altitude of 35,000 feet, at flight Mach numbers of 0.3, 0.8, and 1.2, and at several engine-inlet temperatures,.

"A program was undertaken to determine the J73 turbojet engine compressor stall and surge characteristics and combustor blow-out limits encountered during transient engine operation. Data were obtained in the form of oscillograph traces showing the time history of several engine parameters with changes in engine fuel flow. The data presented in this report are for step and ramp changes in fuel flow at an altitude of 45,000 feet and flight Mach numbers of 0 and 0.8" (p. 1).

Report presenting an investigation to determine if certain combinations of design nozzle-exit velocity distribution and wall contouring can alter the three-dimensional flow characteristics such that nozzle mass flow cannot be predicted by ordinary two-dimensional design techniques. Four nozzle configurations with different wall contours and velocity distributions were tested. Results regarding the comparison of analytical and experimental values of mass flow are provided.

Report presenting an investigation to determine the pressure-recovery and mass-flow characteristics of an inlet configuration designed to be less sensitive to positive angle-of-attack operation than a conventional conical inlet. Results regarding the straight plate, plate swept 47.5, plates swept 42 degrees and 24 degrees, perforated plates, and a general comparison are provided.

Results regarding an investigation to determine the horizontal- and vertical-tail contributions to the static lateral stability of a complete-model swept-wing configuration at high subsonic speeds. Generally, Mach number effects within the range studied and wing effects on the tail contribution were small and the overall trends of the data of the present investigation agreed with those which have been established at low speeds.