Report presenting a free-flight investigation to determine the effect of wing camber and twist at Mach numbers from 1.4 to 2.1 on the lift, drag, and longitudinal stability of a configuration with 52.5 degrees sweptback wing of aspect ratio 3, and inline tail surfaces. Results regarding drag, total normal force and pitching moment, and wash at the horizontal tail are provided.

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.

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.

Memorandum presenting tests conducted with four specific versions of a swept-wing fighter airplane, which indicate that a large reduction in external-drag coefficient was accomplished by redesigning the original configuration. The forebody modifications, which consisted of a smaller canopy, slimmer nose, and sharper inlet lip, reduced the value of the external-drag coefficient from 0.044 to 0.042 at a Mach number of 1.05 and from 0.042 to 0.040 at a Mach number of 1.28.

Report presenting a brief study utilizing pilots to fly a simulator of some longitudinal stability and control problems of an assumed aircraft capable of flights to altitudes essentially out of the atmosphere. The results show that more than the inherent longitudinal damping of the aircraft is necessary to affect a safe flight, particularly during entry into the atmosphere. Results regarding some preliminary considerations, entry, and control in ascent and the ballistics trajectory are provided.

Data were calculated for equivalence ratios of 1 to 3, chamber pressures of 300 and 600 pounds per square inch absolute, and pressure ratios of 1 to 1500. Parameters included are specific impulse, combustion and exit temperature, molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. A correlation is given which permits determination of performance for a wide range of chamber pressures.

Experimental values of damping in roll at zero angle of attack of the Douglas D-558-II research airplane and its components have been obtained at five different Mach numbers. Large effects of Reynolds number, boundary layer, and wing-incidence angle on the damping in roll were obtained.

Memorandum presenting an investigation of a semispan-wing-fuselage model with a 60 degree delta wing with an aspect-ratio-8.2 blunt trailing-edge flap-type control in the 9- by 12-inch blowdown tunnel. Control hinge-moment and effectiveness characteristics were obtained over an angle-of-attack range of plus or minus 10 degrees at control deflections up to 90 degrees.

"Damage to the compressor blades of turbojet engines due to ingestion of foreign objects is a growing problem, the solution of which has been made more difficult by the large percentage of damaging materials that have remained unknown. A rapid emission spectroscopic method was devised to identify the chemical composition of these foreign objects. Results on laboratory-prepared specimens and blades from damaged engines show that the method can be utilized to determine the nature of the ingested foreign objects" (p. 1).

"Measurements of runway roughness were made on runaways at Washington National Airport, New York International Airport, La Guardia Airport, and Logan International Airport, Boston, Mass. This report presents these measurements in the form of tables as well as plots" (p. 1).

Report presenting testing of the performance of a 16-inch tip-diameter rotor designed for impulse operation at the mean radius with a relative Mach number of 1.53 and 114 degrees of turning. Results regarding the rotor with guide vanes, blade-element performance, rotor alone, and modified rotor hub contour are provided.

Memorandum presenting a single J33 combustor and an experimental tubular combustor incorporating a fuel vaporizer as operated with gaseous hydrogen at temperatures slightly above the boiling point of the fuel. Data were obtained to explore possible effects of the fuel temperature on combustor performance and on the control and measurement of fuel flow.

From Introduction: "This report presents the results of an analysis to determine the coolant flow and pressure requirements for a particular return-flow turbine rooter blade design utilizing hydrogen, helium, or air as the coolant."

From Introduction: "This paper presents results of an analysis of shear, bending-moment, and torque loads measured on the vertical tail during rudder-step, rudder-pulse, aileron-roll, and steady-sideslip maneuvers. In order to assess effects of Mach number and altitude, the maneuvers were performed at altitudes of 15,000, 25,000, and 35,000 feet and Mach numbers from 0.49 to 0.82."

Report presenting a friction, wear, and endurance-lift study made with bonded films of mixed oxides containing lead monoxide (PbO) as the main component. The coatings lubricated over the entire temperature range, but were far more effective from 500 to 1250 degrees Fahrenheit than at the lower temperatures. Results regarding the effect of silica additions on coating formation, determination of coating composition, effect of coating thickness on friction and wear, effect of temperature on friction and wear, and endurance properties are provided.

Memorandum presenting the detailed blade-element performance of a five-stage axial-flow compressor for analysis purposes as well as to provide a contribution to the growing body of information necessary for continuing improvement in designing compressors. Results regarding overall and stage performance, blade-element performance, rotor blade-element performance, stator blade-element performance, comparison with design, consideration of boundary-layer blockage effects, and solution of equilibrium equation are provided.

An experimental investigation was carried out to determine the boundary-layer transition characteristics of bodies of revolution with flat and nearly flat faces. Shadowgraphs indicated that the boundary layer remained laminar on the front faces and was turbulent only on the sides. The tests also yielded information on the total drag coefficients and static longitudinal stability of the models.

"An investigation was made at the Langley landing-loads track to obtain data on the coefficient of friction during wheel spin-up. A landing gear was tested at horizontal velocities ranging from 0 to 180 feet per second together with vertical velocities of 7.0 and 9.3 feet per second. The results indicate the effect of forward speed and tire inflation pressure on the coefficient of friction" (p. 1).

Memorandum presenting an investigation of the distortion removal performance and associated total-pressure-loss characteristics of several freely rotating fans, three single-stage and two double-stage freely rotating fans over a range of radial and circumferential distortions of from 0 to 20 percent at inlet annulus Mach numbers from 0.30 to about 0.60. Results regarding individual fan performance, comparison of fan distortion removal characteristics, and fan total-pressure loss and operating speed characteristics are provided.

Memorandum presenting an investigation of the performance of a semielliptical scoop inlet with a two-dimensional flow field at the design Mach number of 2.0. The investigation included a study of the effects of inlet-leading-edge shape and boundary-layer bleed on the pressure recovery and total-pressure distribution. Results regarding performance of inlets without boundary-layer bleed, effect of diverter boundary-layer bleed, effect of leading-edge sweep angle, effect of compression-surface bleed, performance of inlet III-B, tests of buzz suppressors, and tests of rearward-facing control tubes are provided.

"Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors" (p. 1).

"A program was conducted in an altitude facility at the NACA Lewis laboratory to investigate the effects of rapid inlet pressure oscillations on the operation of a current turbo jet engine. These pressure oscillations were approximately sinusoidal in form and were generated to cover a frequency range of 2 to 75 cycles per second and an amplitude range of 10 to 70 percent of the free-stream total pressure. As the oscillation progressed through the compressor, the amplitude was attenuated considerably and a relatively large phase shift (lag) occurred" (p. 1).