Report presenting an investigation to determine the effect of a distributed granular-type roughness on boundary-layer transition at several temperatures. The roughness was not found to introduce any disturbances of significant magnitude to influence transition. Surface cooling did not increase the value of the critical roughness Reynolds number for a distributed granular-type roughness either.

Memorandum presenting experimental force and moment characteristics, including lift-drag ratios, which have been measured and briefly studied for two families of wing-body combinations. One family consisted of essentially arrow wings with half conical bodies. The other consisted of a fineness-ratio-12 body of revolution alone and with flat-plate triangular wings of aspect ratios ranging from 0.375 to 1.8.

Memorandum presenting an investigation in the 4- by 4-foot supersonic pressure tunnel at Mach number 2.01 to determine the longitudinal and lateral stability and control characteristics of a 60 degree delta-wing airplane configuration equipped with a trapezoidal canard control and with wing trailing-edge flap controls. The investigation included studies of the flap control with and without canard surface and studies of the canard control alone and with the flap control.

Report discussing heat data obtained on concave and convex hemispherical nose shapes and hemispherical depressions on a blunted 30-degree nose cone at a variety of Mach numbers. Both hot-jet tests and flight tests were carried out.

"An investigation of the aerodynamic characteristics of several hypersonic missile configurations with various canard controls for an angle-of-attack range from 0 deg to about 28 deg at sideslip angles of about 0 deg and 4 deg at a Mach number of 2.01 has been made in the Langley 4- by 4-foot supersonic pressure tunnel. The configurations tested were a body alone which had a ratio of length to diameter of 10, the body with a 10 deg flare, the body with cruciform fins of 5 deg or 15 deg apex angle, and a flare-stabilized rocket model with a modified Von Karman nose" (p. 1). Various canard surfaces for pitch control only were tested on the body with the 10 deg flare and on the body with both sets of fins. The results indicated that the addition of a flared afterbody or cruciform fins produced configurations which were longitudinally and directionally stable. The body with 5 deg fins should be capable of producing higher normal accelerations than the flared body. A l l of the canard surfaces were effective longitudinal controls which produced net positive increments of normal force and pitching moments which progressively decreased with increasing angle of attack.

From Introduction: "Reported herein are results of the additional direct-stress fatigue tests on: (1) a plate-type specimen with cored centrally located hole and (2) 0.300-inch-diameter round polished specimens with various degrees of porosity."

Report presenting an investigation at Mach number 1.61. to determine the aerodynamic characteristics of a full-span trailing-edge control on a 60 degree delta wing with and without a partial-span spoiler mounted on the wing just ahead of the control. Pressure distribution and hinge-moment measurements were made over a range of angles of attack and control deflection. Results indicated that regions of increase pressure due to flow separation ahead of the control at the larger control deflections and ahead of the spoiler exist.

From Summary: Test results are presented which indicate that the pickup is aerodynamically stable and has accuracy, obtained from a bench calibration, of better than 0.3^o under conditions including acceleration up to 20g in any direction, Mach numbers from 0.5 to 2.8, and dynamic pressures up to at least 65 psi. Equations and curves which can be used to obtain flow direction at the center of gravity of a maneuvering model are presented."

Report presenting an investigation on a full-scale turbojet engine to obtain the altitude-ignition characteristics of three fuels with different ASTM distillation curves. Two of the fuels had a Reid vapor pressure of 2.7 pounds per square inch and the third had a vapor pressure of 1.7 pounds per square inch. Results regarding the effect of fuel volatility and fuel-air ratio on altitude-ignition and propagation characteristics and effect of altitude variation on altitude ignition and propagation characteristics of fuel 53-39 are provided.

Report presenting an experimental investigation to evaluate the internal performance of two types of auxiliary air inlets, a submerged and a scoop inlet, operating within the turbulent boundary layer existing on the bottom of a typical supersonic fighter aircraft afterbody. Results regarding boundary-layer flow and inlet performance are provided.

Memorandum presenting a preliminary evaluation of three types of combustor configuration in a 48-inch-diameter ramjet engine in order to select the one with the greatest promise of efficient and stable combustion at low fuel-air ratios and low combustor-inlet pressures. Results regarding the inlet conditions, performance, and a comparison of the configurations are provided.

"A preliminary evaluation was made of three types of combustor configurations in a 48-inch-diameter ram-jet engine to determine the one with the greatest promise of ultimately achieving efficient and stable combustion at low fuel-air rations and low combustor-inlet pressure. The combustor configurations evaluated include (1) a can-piloted configuration using a gutter-type flame holder with can-type pilot burners for stability, (2) an annular-piloted configuration using a gutter-type flame holder with an annular pilot burner to provide stability and to serve as a flow divider for confined fuel-air mixing, and (3) a can-type configuration using a flow divider for confined fuel-air mixing. No modifications were made to optimize the performance of any configuration" (p. 1).

An investigation was conducted using static capsules fabricated from "L" nickel tubing to determine the effect of temperature level, temperature gradient, and test duration on corrosion and mass transfer by molten sodium hydroxide under free-convection conditions. A base temperature range from 1000 degrees to 1600 degrees F with temperature differences to 500 degrees was studied. The rate of mass transfer was found to be strongly dependent on both temperature level and gradient. The rate shows little tendency to decrease for test durations up to 200 hours, although the concentration of nickel in the melt approaches a limited value after 100 hours.

From Summary: "An analytical investigation was made to determine from component performance characteristics the effect of air bleed at the compressor outlet on the acceleration characteristics of a typical high-pressure-ratio single-spool turbojet engine. Consideration of several operating lines on the compressor performance map with two turbine-inlet temperatures showed that for a minimum acceleration time the turbine-inlet temperature should be the maximum allowable, and the operating line on the compressor map should be as close to the surge region as possible throughout the speed range. Operation along such a line would require a continuously varying bleed area."

"A mixed-flow impeller was designed to give a prescribed blade-surface velocity distribution at mean blade height for a given hub-shroud profile. The blade shape at mean blade height, which was produced by the prescribed velocity distribution, was extended by means of radial lines to form the composite blade shape from hub to shroud. The resulting blade was relatively thick; therefore, it was necessary to retain the inverse blade taper which resulted from extension of the radial lines in order to prevent merging or near merging of the separate blades near the hub. For the first test version of the impeller, designated the MFI-2A, the blade height was arbitrarily made greater than that for the basic impeller (the MFI-2) to allow for viscous effects" (p. 1).

A turbine designed to power a supersonic compressor was investigated with four nozzle settings and one nozzle modification to determine the effect on turbine performance. With the correct ratio of nozzle throat area to rotor effective throat area, the design specific work was obtained near the peak efficiency. From the results of the investigation it was concluded that the convex in sweep of the rotor inner wall and high degree of blade taper caused the turbine configuration to have two aerodynamic disadvantages in comparison with a turbine configuration with an axial inner wall at the rotor outlet and little axial taper.

"An analytical and experimental investigation was made with a J33-A-9 engine to determine the effectiveness of spray cooling as a means of increasing thrust by permitting engine operation at inlet-gas temperatures and speeds above rated. With the assumption of adequate spray cooling at a coolant-to-gas flow ratio of 3 percent, calculations for the sea-level static condition indicated a thrust may be achieved by engine operation at an inlet-gas temperature of 2000 degrees F and an overspeed of 10 percent. Of the water-injection configurations investigated experimentally, those located in the inner ring of the stator diaphragm provided the best cooling at rated engine speed" (p. 1).

Report presenting some measurements of the dynamic lateral stability and lateral and longitudinal trim of a Douglas D-558-II in flights up to a Mach number of 1.87 and an altitude of about 67,000 feet. Testing indicated that the airplane flying in low density air at supersonic speeds had poor dynamic lateral stability, which worsened as the Mach number was increased to 1.85.

Memorandum presenting an investigation of the effects of longitudinal location of the wing and varying body size on the interference characteristics of a 45 degree sweptback wing over a range of Mach numbers and angles of attack. The wing had an aspect ratio of 4, a taper ratio of 0.6, and NACA 65A006 airfoil sections. The results indicated that the zero-lift drag of the wing with interference in the forward position was lower throughout the Mach number range than for the wing with interference in the rearward position.

Report presenting the design, construction, and investigation of an axial-flow compressor inlet stage with a maximum rotor relative inlet Mach number of 1.1. The rotor was designed for a high axial inlet velocity, no inlet guide vanes, and a tip speed of 1000 feet per second. Results regarding the overall performance, rotor inlet conditions, and stator exit conditions are provided.

The report includes a description of the test facilities and methods, crash configuration, layout of crash site, instrumentation, data-recording systems, and the post-crash examination procedure of the fullscale crash-fire investigation, which is part of a comprehensive study of the airplane crash-fire problem.

Models of the Hermes A-3B missile were tested in the Ames supersonic free-flight wind tunnel to determine the static-longitudinal-stability characteristics at a Mach number of 5.0 and a Reynolds number based on body length of 10 million. The results indicated that the model center of pressure was 45.3 percent of the body length aft of the nose and the lift-curve slope based on body frontal area was 0.064 per degree. Estimates indicated that the effect on these characteristics of aeroelastic twisting of the model fins was small but important if a precise location of center of pressure is required. A comparison of the test results with predictions based on available theory showed that the theory was useful only for rough estimates, The drag coefficient at zero lift, based on body frontal area, was found to be 0.155.

Report presenting estimates of downwash angles determined from an analysis of pressures measured on the horizontal tail of a supersonic aircraft configuration. The results are compared to an approximate application of linearized theory and force-test results for the same configuration.

Report presenting an investigation of the effect of separation vortex flow on the downwash, sidewash, and wake characteristics behind a 47.5 degree sweptback wing with symmetrical circular-arc airfoil sections. Three configurations were investigated: the basic wing, the wing with full-span drooped-nose flaps deflected 40 degrees, and the wing with semispan plain flaps and full-span drooped-nose flaps deflected 40 degrees. Results regarding the air-stream surveys, distribution of vorticity, average values of downwash and dynamic-pressure ratio, and correlation with theory are provided.