From Summary: "A preliminary investigation was conducted to determine the altitude performance of pentaborane and a blend fuel of 64.2 percent pentaborane in JP-4 fuel in a tubular combustor of a current production type turbojet engine. Deposits on the liner barrel and dome were virtually nonexistent. The combustion efficiency of 90 to 94 percent for pentaborane and approximately 90 percent for the blend. Outlet temperature profiles were marginal; however, the developed combustor provided lower pressure losses than obtained in conventional combustors."

Report presenting an investigation in the 7- by 10-foot tunnel to determine the effect of aspect ratio on the aerodynamic characteristics in pitch of wing-fuselage combinations with 45 degrees sweepback at the quarter-chord line and 0.6 taper ratio at high subsonic speeds. Results regarding the lift characteristics, pitching-moment characteristics, drag characteristics, and lift-drag ratios are provided.

Report presenting a compilation of free-spinning model results of investigations of the spin and recovery characteristics of 12 flying-wing and unconventional-type designs. Dimensional data, mass data, and three-view drawings of the models that correspond to each of the 12 airplane designs are provided. Results regarding erect spins, inverted spins, and spin-recovery parachutes are also provided.

Report presenting the performance of the turbine component of a Python turbine-propeller engine with four tail-pipe configurations determined over a range of altitudes, engine speeds, and fuel flows. Results regarding the corrected turbine speed, corrected turbine enthalpy drop, pressure ratio, and efficiency were established on a turbine-characteristic plot.

Report presenting an investigation in the altitude wind tunnel to determine the performance of the Python turbine-propeller engine. Compressor-performance data were obtained for a range of simulated altitudes at a single cowl-inlet ram pressure ratio while the engine was run over its full operable range of engine speeds. Results regarding the deterioration of compressor performance, compressor performance, including efficiency and performance maps, and velocity and static-pressure profiles are provided.

Memorandum presenting an analysis made to estimate the performance of a direct-air nuclear turbojet-powered airplane with a split-flow reactor and a separated-type shield. The analysis was for flight Mach numbers of 0.9 and 1.5 and covered a range of altitudes, reactor-wall temperatures, turbine-inlet temperatures, compressor pressure ratios, and airplane lift-drag ratios.

Report presenting an investigation to determine the longitudinal control characteristics of deflectable wing-tip elevators on a low-aspect-ratio, untapered, 45 degree sweptback semispan wing. The elevators had triangular and parallelogrammic plan forms and flat-plate profiles. Results include the elevator effectiveness and comparison of the elevator effectiveness of deflectable wing-tip controls and trailing-edge flap-type controls.

"The NACA has conducted an investigation to determine the performance that can be obtained from a multistage axial-flow compressor based on airfoil research. A theory was developed; an eight-stage axial-flow compressor was designed, constructed, and tested. The performance of the compressor was determined for speeds from 5000 to 14,000 r.p.m with varying air flow at each speed" (p. 81).

A method of constructing fins of nearly optimum proportions has been developed by the NACA to the point where a cylinder has been manufactured and tested. Data were obtained on cylinder temperature for a wide range of inlet-manifold pressures, engine speeds, and cooling-pressure differences. The results indicate that an improvement of 40 percent in the outside-wall heat-transfer coefficient could be realized on the present NACA cylinder by providing a thermal bond equivalent to that of an integral fin-cylinder wall combination between the preformed fins and the cast cylinder wall" (p. 107).

"Theoretical investigations have shown that, because air is compressible, the pressure-drop requirements for cooling an air-cooled engine will be much greater at high altitudes and high speeds than at sea level and low speeds. Tests were conducted by the NACA to obtain some experimental confirmation of the effect of air compressibility on cooling and pressure loss of a baffled cylinder barrel and to evaluate various methods of analysis. The results reported in the present paper are regarded as preliminary to tests on single-cylinder and multicylinder engines. Tests were conducted over a wide range of air flows and density altitudes" (p. 1).

Preignition-limited performance data were obtained on a supercharged CFR engine at two sets of operating conditions over a wide range of fuel-air ratios to determine the preignition characteristics for the following five fuels: s-3 reference fuel, s-3 plus 4 ml. tel per gallon, afd-33(140-p), benzene, and diisobutylene. Maximum thermal-plug temperatures at constant intake-air pressures were also determined to correlate the preignition characteristics of each fuel with its ability to increase general engine-temperature levels. Additional runs were made to compare the preignition-limited performance of triptane, triptane plus 4 ml. tel per gallon, and an-f-28r fuel.

The effect of internal cooling on the knock-limited performance of an-f-28 fuel was investigated in a CFR engine, and the following internal coolants were used: (1) water, (2), methyl alcohol-water mixture, (3) ammonia-methyl alcohol-water mixture, (4) monomethylamine-water mixture, (5) dimethylamine-water mixture, and (6) trimethylamine-water mixture. Tests were run at inlet-air temperatures of 150 degrees and 250 degrees F. to indicate the temperature sensitivity of the internal-coolant solutions.

"Several designs of nitrided-steel piston rings were performance-tested under variable conditions of output. The necessity of good surface finish and conformity of the ring to the bore was indicated in the preliminary tests. Nitrided-steel rings of the same dimensions as cast-iron rings operating on the original piston were unsatisfactory, and the final design was a lighter, rectangular, thin-face-width ring used on a piston having a maximum cross-head area and a revised skirt shape. Results were obtained from single-cylinder and multicylinder engine runs" (p. 1).

From Introduction: "The purpose of this analysis was to investigate the use of a function of intake-manifold pressure, exhaust back pressure, intake manifold temperature, and engine speed in place of a venturi as a means of measuring engine air consumption and to determine if this function is suitable for automatic mixture control."

"An experimental investigation of rectangular exhaust-gas ejector pumps was conducted to provide data that would serve as a guide to the design of ejector applications for aircraft engines with marginal cooling. The pumping characteristics of rectangular ejectors actuated by the exhaust of a single-cylinder aircraft engine were determined for a range of ejector mixing-section area from 20 to 50 square inches, over-all length from 12 to 42 inches, aspect ratio from 1 to 5, diffusing exit area from 20 to 81 square inches, and exhaust-nozzle aspect ratio from 1 to 42" (p. 161).

From Summary: "A ceramic consisting nominally of 75 percent molybdenum disilicide and 25 percent alumium oxide was investigated to determine its thermal shock characteristics and high-temperature strength properties. In a rim-quench thermal shock evaluation, the material was found to be superior to pure molybdenum disilicide. In a simulated altitude blow-out test, the material withstood an average of 2.5 cycles when quenched from 1800 degrees Fahrenheit."

Report presenting a flight investigation to determine horizontal-tail loads at transonic speeds of the Bell X-1. The tests were made through the transonic region and at high lift coefficients. Results regarding the time histories, tail load data, aerodynamic-center positions, and potential load discrepancy explanations are presented.

Report presenting zero-lift drag data for a series of fin-stabilized bodies differing only in afterbody shape. Three series of afterbodies with varying ratios of base radium to maximum radius were combined with a fineness-ratio-7.13 parabolic nose. Results regarding the total drag, drag breakdown, comparison of experimental and theoretical afterbody pressure drag, and afterbody design criteria are provided.

Report presenting an investigation of the lift, drag, pitching-moment, and downwash characteristics of a triangular wing of aspect ratio 3 with an NACA 63A006 section. Downwash was measured by using an all-movable horizontal tail at five locations. Results regarding basic wing characteristics and stability contribution of the tail are also provided.

Report presenting an investigation to determine the low-speed characteristics of a variable-sweep airplane model with a twisted and cambered wing. The wing was designed to have a uniform load distribution at a Mach number 1.10 and a lift coefficient of 0.25 and 50 degrees sweep. Results regarding lift characteristics, drag characteristics, lift-drag ratio, pitching-moment characteristics, dynamic pressure and downwash at the tail, effect of flaps, and lateral stability characteristics are provided.

Report presenting an investigation in the 4- by 4-foot supersonic pressure tunnel at Mach number 2.01 to determine the effects of sweep angle and thickness ratio on the static lateral stability characteristics of a series of wings with taper ratio 0.2 and aspect ratio 3.5. The results are presented with little analysis.

"In recent months interest has increased in the possible use of distillate and residual fuel oils as fuels for commercial gas-turbine aircraft. However, the use of such fuels entails the solution of many problems pertaining to fuel physical properties and combustion characteristics. This report reviews some of these problems and discusses the status of current knowledge in relation to their solution" (p. 1).

Report presenting testing of a model of a Navy Bureau of Aeronautics design in Langley tank no. 2 to determine the calm-water take-off and rough-water landing characteristics of the design with particular attention paid to the take-off resistance and landing accelerations. Results regarding resistance, take-off stability, landing stability, spray, and landing acceleration are provided.

Report reporting measurements of pressure distributions obtained in a flight test of a fin-stabilized parabolic body of revolution with a sharply convergent afterbody. Pressures were measured at eight longitudinal stations on the body and over a range of Mach numbers. Results regarding basic data, the effect of sting, the variation of pressure coefficient with Mach number, pressure distributions, velocity distribution and discussion of flow phenomena, and drag are presented.