Performance characteristics of a 16-inch annular-type combustor installed in a full-scale engine using gaseous-hydrogen fuel were obtained at simulated altitudes from 66,000 to 86,000 feet at a flight Mach number of 0.8. Combustion efficiencies of 86 percent were obtained at 86,000 feet (combustor pressure, 420 lb/sq ft abs). Combustor blowout was not encountered during the investigation.

Report presenting an investigation in the altitude test chamber to evaluate the performance of a 16-inch-long combustor designed for use with gaseous-hydrogen fuel. The investigation covered a range of combustor pressures and simulated altitudes at Mach number 0.8. Results regarding operational characteristics, combustor performance, and temperature profiles are provided.

An investigation of the effect of extremely high altitude operation on the performance and operating characteristics of two turbojet engines using gaseous hydrogen and JP-4 fuels. At extremely high altitude conditions, engine performance was significantly poorer than at low altitudes. The majority of performance losses were due to the compressor because of low Reynolds number and the combustor because of low combustion efficiency.

An investigation was conducted in the NACA Lewis altitude wind tunnel to determine the effect of compressor-outlet bleedoff on the performance of an axial-flow turbojet engine equipped with a variable-area exhaust nozzle. Results presented indicate the effect of compressor-outlet bleedoff on performance at altitudes of 25,000 and 40,000 feet and a flight Mach number of 0.53. Variation of performance with bleedoff flow is indicated for operation with fixed- and variable-area exhaust nozzles. Temperature and pressure losses through the bleedoff ducting system are also discussed.

Report of an investigation into the effect of leading-edge radius on the high-speed aerodynamic characteristics of a symmetrical, 9-percent-thick airfoil using three different airfoils at several Mach and Reynolds numbers. Information about the normal-force coefficients, drag coefficients, moment coefficients about the quarter chord, pressure-distribution diagrams, Schileren photographs, and a comparison with data obtained in other tunnels is included.

Report presenting results from a tank investigation of a powered dynamic model of the Grumman JRF-5 airplane fitted with a single hydro-ski and extended afterbody. Results indicated that the afterbody extension in place of a tail ski reduced the maximum resistance by 10 percent. Removal of the wing-tip skids gave a further reduction in maximum resistance of 3 percent.

Report presenting results of pressure-distribution measurements at low speed on a wing-fuselage combination with a wing of W plan form with aspect ratio 6, 45 degrees sweep, taper ratio 0.6, and an NACA 65A009 airfoil section placed parallel to the plane of symmetry. Results regarding aerodynamic characteristics, pressure distribution on the wing, pressure distribution on the fuselage, aerodynamic section characteristics, and stall patterns are provided.

Report presenting an investigation by the wing-flow method to determine the snaking characteristics of a scale partial-model of a research airplane. Additionally, the snaking characteristics of 11 modified configurations of the research-airplane model were determined as well as a model that had a greater fineness ratio.

Report presenting an investigation in the free-flight tunnel to determine the dynamic lateral stability and control characteristics of a high-aspect-ratio bomber model with self-supporting, free-floating fuel tanks attached to the wing tips. The results showed that the flight behavior of the configuration could be made satisfactory when sufficient restoring moment was supplied. Results regarding the flight behavior, rolling motions, motions of the bomber and attached tanks, and flight records for different gearing ratios are provided.

Memorandum presenting an investigation of the leading-edge extensions, a trailing-edge extension, and a fence on the static longitudinal stability of a wing-fuselage-tail combination with a wing with 35 degrees of sweepback and an aspect ratio of 4.5. The investigation involved the use of force measurements and tuft studies on stall progression. Results regarding a model with unmodified wing, the effect of leading-edge modifications, the effects of the fence, a summary of the effects of compressibility, and some remarks on flow separation are provided.

Memorandum presenting an investigation of operation of a single-spool turbojet engine with variable exhaust-nozzle area over a range of altitudes at a constant flight Mach number. Data was obtained by subjecting the engine to approximate step disturbances in the independent variables, and the information necessary to effect a linearized first-order description of the engine's dynamic operation was obtained. Results regarding rotational speed, compressor-discharge pressure, turbine-discharge pressure, net thrust, and turbine-discharge temperature are provided.

Report presents an investigation to determine the transonic aerodynamic characteristics of W-plan-form wing of aspect ratio 3.5, taper ratio 0.25, and 60 degree 48 minute panel sweep by using a small reflection plane over which high subsonic and low subsonic Mach numbers are obtained. Comparisons are made between sweptback-wing data to determine the effectiveness of the W-plan modifications.

Characteristics of NACA submerged duct entries and wing leading-edge inlets designed for a 1/4 scale flow model of a fighter-type airplane powered by a jet engine in the fuselage are presented.

Two current turbojet engines were operated with gaseous-hydrogen and JP-4 fuels at very high altitudes and a simulated Mach number of 0.8. With gaseous hydrogen as the fuel stable operation was obtained at altitudes up to the facility limit of about 90,000 feet and the specific fuel consumption was only 40 percent of that with JP-4 fuel. With JP-4 as the fuel combustion was unstable at altitudes above 60,000 to 65,000 feet and blowout limits were reached at 75,000 to 80,000 feet. Over-all performance, component efficiencies, and operating range were reduced considerable at very high altitudes with both fuels.

An investigation is in progress at the Langley Laboratory of the NACA to explore the possibilities of axial-flow compressors operating with supersonic velocities relative to the blade rows. The first phase of this investigation, a study of supersonic diffusers, has been reported. The second phase, an analysis of supersonic compressors, has also been reported. Preliminary calculations have shown that very high pressure ratios across a stage, together with somewhat increased mass flows, are possible with compressors which decelerate air through the speed of sound in their rotor blading. These performance characteristics are desirable in compressors for aircraft jet propulsion units, gas turbines, or superchargers. The third phase, presented here, is a preliminary experimental investigation of a supersonic compressor designed to produce a high pressure ratio in a single stage.