From Introduction: "In this report, altitude, airplane velocity, exhaust-nozzle area, and turbine-inlet temperature were chosen and corresponding engine speed, fuel flow, and throttle pressure ratio were found. A detailed discussion of the engine analysis and the calculations used are presented in Appendix A. The derivations of the principal equations used are presented in appendix B."

Report presenting a flight investigation in natural icing conditions to determine the effect of ice formations on the performance of an axial-flow turbojet engine. Results regarding the tail-pipe temperature, engine jet thrust, and characteristics of ice formation are provided. No general conclusions can be reached from the data because the icing condition was relatively light.

"Low-speed wind-tunnel tests were made of a triangular wing with two arrangements of twin vertical fins. With these twin fins mounted either symmetrically or wholly above the wing chord plane, approximately constant effectiveness in proportion to the fin size was obtained throughout the angle-of-attack range. The addition of these fins, however, resulted in a decrease in maximum lift and a reduction in static longitudinal stability at lift coefficients above 0.4" (p. 1).

Report presenting the performance characteristics of the I-40 jet propulsion engine and an analysis for a range of altitudes and ram pressure ratios. Results regarding the component efficiencies, engine performance, generalizing factors, and windmilling drag characteristics are provided.

Report presenting the performance characteristics of the I-40 jet-propulsion engine for a range of altitudes and ram pressure ratios. Results regarding component efficiencies, engine performance, generalizing factors, and windmilling drag characteristics are provided.

Report presenting the results of an estimation of the limit forces on airfoils at supersonic speeds. The limit pressure coefficient attainable on an airfoil is shown to be about 70 percent of the pressure coefficient for a vacuum over a wide range of Mach numbers.

Report presenting the use of Von Karman's two-dimensional transonic similarity law as applied to drag data of three different thickness wings with NACA 65-series sections and aspect ratios of 7.6 over a range of Mach numbers. The correlation was found to be satisfactory, with results better in the subsonic range than in the supersonic range.

Report presenting an investigation of burner liners in a turbojet engine to determine the principal factors limiting the burner-liner service life. The investigation covered a range of engine speeds and testing was conducted to determine whether bare, ceramic-coated, or shielded thermocouples would give the most correct temperature readings.

Report presenting measurements of the chordwise pressure distributions over the 8-percent-thick wing of the XS-1 research airplane at a section near the midspan of the left wing. Data are presented for a Mach number range and a normal-force coefficient of about 0.33. Results regarding the upper-surface pressure distribution, lower-surface pressure distributions, and total section loads are provided.

Memorandum presenting a description of an axial-discharge mixed-flow compressor, which is especially adapted for jet engines because of the large mass flow per unit frontal area. The best impeller was selected on the basis of the maximum air-flow capacity, which was 19.6 pounds per second for a 14-inch-diameter impeller with a tip speed of 1480 feet per second and a pressure ratio of 3.5.

Report presenting an analysis of an axial-discharge mixed-flow compressor designed to combine the compactness, reliability, and wide operating range of a mixed-flow compressor with the high flow capacity per unit of frontal area that characterizes the axial-flow compressor. The primary objective of this design was to maximize the flow capacity for a predetermined pressure ratio. Results regarding the flow distribution at the impeller inlet, flow distribution at the impeller outlet, and impeller performance are provided.

"A flight investigation was conducted with a fighter-type airplane to determine the loss in rolling effectiveness due to wing twist, and the results are compared with the loss in rolling effectiveness calculated by use of a simplified method. The computed rolling velocities showed good agreement with experimentally determined rolling velocities throughout the Mach number range covered by the flights. Extrapolation of the flight data indicated a Mach number for reversal which agreed well with that obtained from calculations" (p. 1).

Memorandum presenting an investigation of a model with a ramp that produced an oblique shock wave in front of an annular duct inlet, which was tested at Mach numbers between 1.36 and 2.01 to determine the effect of reducing the entrance Mach number on the total-pressure recovery after diffusion. The results showed that the maximum total-pressure recovery of the model investigated was considerably improved through the addition of a ramp.

Report presenting an investigation to determine the performance of a Bumblebee 18-inch ram jet with a rake-type flame holder was conducted using kerosene and propylene oxide as fuels. Experiments occurred at a range of pressure altitudes and ram-pressure ratios over the operable range of fuel-air ratios. Use of propylene oxide was found to improve the stability of combustion, increased the operable range of fuel-air ratios, and allowed ignition at high values of combustion-chamber-inlet velocities.

Report presenting testing of several delta wing-aileron configurations as part of a general investigation of their rolling effectiveness characteristics at transonic and supersonic speeds. Results regarding wing-aileron rolling effectiveness and drag are provided.

Report presenting an investigation to determine the characteristics in yaw of a trapezoidal wing of aspect ratio 4 with circular-arc airfoil sections and 0.20-chord rear and drooped-nose flaps. Measurements of lift, lateral force, rolling moment, and yawing moment of four configurations were made for a range of angles of attack and yaw angles.

"A two-dimensional wedge diffuser, designed to be used with a ducted-airfoil ram jet, was tested on a rocket-powered test vehicle up to a Mach number of 1.45. Orifice plates and a choking section behind the diffuser exit simulated burning in a combustion chamber by providing the proper pressure drop. Results obtained from these tests showed there was no abrupt changes in mass flow and pressure recovery as the model velocity increased through the transonic region" (p. 1).

Report presenting the results of an investigation in the altitude wind tunnel to determine the performance of a Bumblebee 18-inch ram jet equipped with a can-type flame holder. Results regarding the ramjet combustion performance and ramjet overall performance are provided.

From Introduction: "The particular blade design, the tests of which are described in the present paper, was one necessary to the investigation of the effect of differences in blade-thickness ratios."

"Finally the work of the report is applied exclusively to consideration of the turbine-propeller power plant because previous performance analyses of the various gas-turbine propulsion systems combine with recent improvements in propeller design indicate that it is this combination which, will give the best over-all performance in the speed ranges considered" (p. 5).

From Summary: "Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle."

"An analysis of the performance of the types A, B, and C combustion chambers of the 4000-pound-thrust axial-flow turbojet engine is presented. The data were obtained from investigations of the complete engine over a range of pressure altitudes from 5000 to 40,000 feet and ram pressure ratios from 1.00 to 1.86. The combustion-chamber pressure losses, the effect of the losses on cycle efficiency, and the combustion efficiency are discussed" (p. 1).

"The J33-A-23 compressor with a 34-blade impeller was operated at ambient inlet temperature and an inlet pressure of 14 inches mercury absolute over a range of equivalent impeller speeds from 6000 to 11,750 rpm. Additional runs at equivalent speeds of 7,000, 10,000, and 11,750 rpm and ambient inlet temperature were made at inlet pressures of 5 and 10 inches mercury absolute. The results of this investigation are compared with those of the J33-A-23 compressor with a 17-blade impeller" (p. 1).

"As part of an investigation of the application of nuclear energy to various types of power plants for aircraft, calculations have been made to determine the effect of several operating conditions on the performance of condensers for mercury-turbine power plants. The analysis covered 8 range of turbine-outlet pressures from 1 to 200 pounds per square inch absolute, turbine-inlet pressures from 300 to 700 pounds per square inch absolute,and a range of condenser cooling-air pressure drops, airplane flight speeds, and altitudes. The maximum load-carrying capacity (available for the nuclear reactor, working fluid, and cargo) of a mercury-turbine powered aircraft would be about half the gross weight of the airplane at a flight speed of 509 miles per hour and an altitude of 30,000 feet" (p. 1).