Report presenting an examination of the use of an axisymmetric variable-geometry technique for achieving good supersonic inlet performance over a wide range of Mach numbers. An inlet with a telescoping spike was designed and experimentally evaluated. Results regarding the spike survey study and variable-geometry-inlet study are provided.

Report presenting an experimental investigation at Mach number 1.90 to determine the overall performance capabilities of axisymmetric two-cone and isentropic nose inlets in terms of total-pressure recovery, mass flow, and external drags. At zero angle of attack, the external drag was separated into their components of cowl pressure, friction, and additive drags. Results regarding total-pressure recovery and mass-flow characteristics, total-pressure profiles at diffuser exit, cowl-pressure distributions and drag, zero-angle-of-attack component breakdown of external drag, and overall performance comparison are provided.

Report presenting a centrifugal impeller with a modified design procedure to reduce the velocity gradients existing along the hub from inlet to outlet. The modified impeller had better performance characteristics than the original at all speeds tested.

Report presenting an analysis made to permit a comparative evaluation of the turbine rotor cooling-air requirements with convection-cooled blades for four combustor-outlet radial gas-temperature profiles: a uniform profile, a profile representable by a complete cycle of the cosine wave using cooler gas layers near the turbine casing, a profile representable by one-half cycle of the cosine wave using cooler gas layers near the stator inner ring, and an optimum profile resulting in the absolute minimum cooling-air requirements.

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 an investigation at transonic speeds in the high speed tunnel to determine the rolling-effectiveness characteristics of several aileron arrangements for use on a 42 degree sweptback wing. Results regarding the rolling-moment coefficients, variation of coefficient with Mach number, outboard cusp aileron, and use of the leading-edge flaps are provided.

Memorandum presenting a method for computing the isentropic and actual enthalpy change between the inlet and outlet of a centrifugal-flow compressor when water injection is used. The method of calculation is based on fluid property values given in steam and air tables.

Report presenting a method for calculating the path of a jettisonable nose section at any speed by means of successive approximations using static aerodynamic characteristics. Conditions for consideration, including the effect of stabilizing fins, are included in the calculations.

Report presenting an investigation of the performance of a pair of tubular turbojet combustors interconnected by a 3-inch-inside-diameter ceramic-lined pilot chamber to determine the utility of the pilot chamber for improvement of high-altitude combustion efficiency, lean limits, and minimum engine speed. Results regarding the combustion efficiency and lean-limit blow-out, ignition limits, and altitude operating limits are provided.

Performance characteristics of conical jet nozzles were determined in an investigation covering a range of pressure ratios from 1.0 to 2.8, cone half-angles from 5 degrees to 90 degrees, and outlet-inlet diameter ratios from 0.50 to 0.91. All nozzles investigated had an inlet diameter of 5 inches. The flow coefficients of the conical nozzles investigated were dependent on the cone half-angle, outlet-inlet diameter ratio, and pressure ratio. The velocity coefficients were essentially constant at pressure ratios below the critical. For increasing pressures above critical pressure ratio, there was a small decrease in velocity coefficient that was dependent on pressure ratio and independent of cone half-angle and outlet-inlet diameter ratio. Therefore the variation in performance (air flow and thrust) of several nozzles, selected for the same performance at a particular design condition, was proportional to the ratio of their flow coefficients.

Report presenting free-falling recoverable model tests conducted at transonic speeds on a model with a triangular wing of aspect ratio 2 and a 45 degree swept tail in the chord plane of the wing. Static- and dynamic-longitudinal-stability data for the complete model, force and moment data for the major components of the model, and load distributions over the fuselage of the model were evaluated at several angles of attack.

Report presenting a preliminary investigation of a variable-area, variable-internal-contraction air inlet at a range of Mach numbers. The total pressure recovery after supersonic compression, total pressure recovery at the exit of the subsonic diffuser, and the pressure distribution along the diffuser walls were measured.

Six large 24s-t3 aluminum-alloy-sheet specimens containing various notches or fillets were tested in tension to determine their stress concentration factors in both the elastic and plastic ranges. The elastic stress concentration factors were found to be slightly higher than those calculated by Neuber's method and those obtained photoelastically by Frocht. The results showed further that the stress concentration factor decreases as strains at the discontinuity enter the plastic range. A generalization of Stowell's relation for the plastic stress concentration factor at a circular hole in an infinite plate was applied to the specimen shapes tested and gave good agreement with test results.

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.

"Results of an investigation by the free-fall method are presented herein for a configuration having a body of revolution of fineness ratio 12 and 45 degrees sweptback wing mounted aft of the maximum diameter of the body. The fillets were designed to provide large increases in the sweep of the leading edge and the line of maximum thickness as the wing root was approached. Comparison of these results with those for the same configuration without fillets shows that the addition of wing fillets increased the total drag of the configuration by about 35 percent at Mach numbers near 1.0 and about 15 percent at Mach numbers near 1.2" (p. 1).

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.

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.

From Summary: "Flight tests have been conducted on rocket-propelled models of an airplane configuration incorporating a sweptback wing with inverse taper to investigate the drag, stability, and control characteristics at transonic and supersonic speeds. The models were tested with a conventional tail arrangement in the Mach number range from 0.55 to 1.2. In addition to the various aerodynamic parameters obtained, the flying qualities were computed for a full-scale airplane with the center-of-gravity location at 18 percent of the mean aerodynamic chord. Also, included in this investigation are drag measurements made on relatively simple fixed-control models tested with both conventional and V-tail arrangements."

This report presents the results of an investigation conducted to determine the interrelation of the constituents of the exhaust gases of internal-combustion engines and the effect of engine performance on these relations. Six single-cylinder, liquid-cooled tests engines and one 9-cylinder radial air-cooled engine were tested. Various types of combustion chambers were used and the engines were operated at compression ratios from 5.1 to 7.0 using spark ignition and from 13.5 to 15.6 using compression ignition. The investigation covered a range of engine speeds from 1,500 to 2,100 r.p.m.