Shock-wave equations have been evaluated for a range of Mach number in front of the shock from 1.05 to 4.0. Mach number behind the shock, pressure ratio, derivation of flow, and angle of shock are presented on charts. Values are also included for density ratio and change in entropy.

From Summary: "Axial fatigue tests at zero mean stress have been made on 0.032- and 0.064-inch 24S-T and 0.032-inch 75S-T sheet-metal specimens 1/4, 1/2, 1, and 2 inches wide without a hole and with central holes giving a range of hole diameter D to specimen width W from 0.01 to 0.95. No systematic difference was noted between the results for the 0.032-inch and the 0.064-inch specimens although the latter seemed the more consistent. In general the fatigue strength based on the minimum section dropped sharply as the ration D/W was increased from zero to about 0.25. The plain specimens showed quite a pronounced decrease in fatigue strength with increasing width. The holed specimens showed only slight and rather inconclusive evidence of this size effect. The fatigue stress-concentration factor was higher for 75S-T than for 24S-T alloy. Evidence was found that a very small hole would not cause any reduction in fatigue strength."

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).

Contains theoretical and experimental analysis of circular inlets having a central body at Mach numbers of 3.30, 2.75, and 2.45. The inlets have been designed in order to have low drag and high pressure recovery. The pressure recoveries obtained are of the same order of magnitude as those previously obtained by inlets having very large external drag.

"Data obtained from an extensive investigation of the cooling characteristics of four multicylinder, liquid-cooled engines have been analyzed and a correlation of both the cylinder-head temperatures and the coolant heat rejections with the primary engine and coolant variables was obtained. The method of correlation was previously developed by the NACA from an analysis of the cooling processes involved in a liquid-cooled-engine cylinder and is based on the theory of nonboiling, forced-convection heat transfer. The data correlated included engine power outputs from 275 to 1860 brake horsepower; coolant flows from 50 to 320 gallons per minute; coolants varying in composition from 100 percent water to 97 percent ethylene glycol and 3 percent water; and ranges of engine speed, manifold pressure, carburetor-air temperature, fuel-air ratio, exhaust-gas pressure, ignition timing, and coolant temperature" (p. 207).