"Tests have been conducted in the Langley 8-foot transonic tunnel on a 0.04956-scale model of the Convair F-102A airplane which employed an indented and extended fuselage, cambered wing leading edges, and deflected wing tips. Force and moment characteristics were obtained for Mach numbers from 0.60 to 1.135 at angles of attack up to 20 degrees. In addition, tests were made over a limited angle-of-attack range to determine the effects of the cambered leading edges, deflected tips, and a nose section with a smooth area distribution" (p. 1).

"Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.86, a Reynolds number of 129,000 based on diameter, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and with a simple modification of the Newtonian flow theory. The comparison of experimental results shows that either theory gives adequate general aerodynamic characteristics but that the modified Newtonian theory gives a more accurate prediction of the pressure distribution" (p. 1).

"A summary and analysis have been made of the results of various investigations to determine the aerodynamic characteristics of a supersonic aircraft configuration. The configuration has a wing with 40 degree sweepback at the quarter-chord line, aspect ratio 4, taper ratio 0.5, and 10-percent-thick circular-arc sections normal to the quarter-chord line. Experimental data were available for a Mach number range from 0.16 to 2.32. Results obtained from wing-flow, rocket-model, transonic-bump, and tunnel tests are presented and, where possible, are supplemented by empirical and theoretical calculations" (p. 1).

"Available experimental two-dimensional cascade data for conventional compressor blade sections are correlated at a reference incidence angle in the region of minimum loss. Variations of reference incidence angle, total-pressure loss, and deviation angle with cascade geometry, inlet Mach number, and Reynolds number are investigated. From the analysis and the correlations of the available data, rules and relations are evolved for the prediction of blade-profile performance" (p. 1).

From Introduction: "The primary purpose of the wing instrumentation was to obtain data for use in design of NACA research missiles; however, it is believed that the wing is representative of typical supersonic designs and that the data will be of general use and interest. Only the wing data are reported herein."

From Introduction: "The importance of blunt noses as a means of reducing the heat transfer to high velocity missiles has recently received much publicity. The question of just what blunt shape is best is still moot, and it is the purpose of this paper to present and examine some recent experimental results which may throw some light on this problem."

From Introduction: "The importance of blunt noses as a means of reducing the heat transfer to high velocity missiles has recently received much publicity. The question of just what blunt shape is best is still moot, and it is the purpose of this paper to present and examine some recent experimental results which may throw some light on the problem."

From Introduction: "Data are presented to show the effects of tail-pipe fuel-air ratio, altitude, and flight Mach number on tail-pipe-burner performance at rated engine speed and approximately constant turbine-outlet temperature. Operational characteristics of the tail-pipe burner and variable-area exhaust nozzle are also reported."

Report presenting an investigation of the estimated minimum cooling requirements and related performance of turbojet engines equipped with high-performance single-stage turbines with air-cooled corrugated-insert blades over a range of turbine-inlet temperature, tip speed, and hub-tip radius ratio for Mach number 2 at 50,000 feet. The effects of stress-ratio factor, flight Mach number, altitude, turbine rotor impeller efficiency, and outside heat-transfer coefficient on cooling requirements were also investigated.

From Introduction: "Analyses of heat transfer for liquid metals flowing turbulently in smooth tubes (low Pradtl numbers) are given in reference 1 and 2. The predicted heat-transfer coefficients from these analyses are considerably higher than those determined in the experimental heat-transfer investigations for mercury and lead bismuth given in references 2 to 4. If the analytical and experimental results for flow of gases are compared at low Peclet numbers, they are also found to be in disagreement (references 5 and 6)."

"Flight icing-rate data obtained in a dense and abnormally deep supercooled stratiform cloud system indicated the existence of liquid-water contents generally exceeding values in amount and extent previously reported over the midwestern sections of the United States. Additional information obtained during descent through a part of the cloud system indicated liquid-water contents that significantly exceeded theoretical values, especially near the middle of the cloud layer. The growth of cloud droplets to sizes that resulted in sedimentation from the upper portions of the cloud is considered to be a possible cause of the high water contents near the center of the cloud layer" (p. 1).

Memorandum presenting an investigation of the theoretical performance of a two-stage ballistic rocket missile with a centerbody and two parallel boosters for JP4-oxygen and ammonia-fluorine propellants. The missile using the ammonia-fluorine propellant weighs about one-half as much as a missile using JP4-oxygen. Results regarding the optimum missile and effect of variation of design parameters are provided.

From Introduction: "The investigation reported herein is a continuation of a ram-jet-combustor design program being conducted at the NACA Lewis laboratory. The purpose of this broad program is to establish basic design criteria for combustors operating over wide range of fuel-air ratio with low pressure losses and high combustion efficiency, and to utilize these design criteria in the development of practical ram-jet combustors."

Analytical methods are presented for the determination of local values of outside and inside heat-transfer coefficients and effective gas temperatures by use of turbine-blade-temperature measurements. The methods are derived for a number of configurations that can be applied to typical cooled-turbine-blade shapes as well as to other types of heat-transfer apparatus.

"An analysis was made at flight Mach number of 0.9 to estimate performance of nuclear-energy-powered turbojet engine and optimum engine operating conditions and to determine gross weight and load-carrying capacity of airplane powered by such an engine. The size of airplane required to carry disposable load of 20,000 pounds was found to vary from approximately 300,000 to 900,000 pounds depending on assumptions. For a reactor tube-wall mean temperature of 2500 degrees R, turbine-inlet temperature of 2000 degrees R, reactor-free-flow-area ratio of 0.33, reactor-shielding-material specific gravity of 6.0, shielding thickness of 3.0 feet, and altitude of 30,000 feet, the airplane gross weight required to carry a 20,000 payload is 545,000 pounds" (p. 1).

From Summary: "Signal Corps wind equipment AN/GMQ-1 consisting of a 3-cup anemometer and wind vane was calibrated for wind velocities from 1 to 200 miles per hour. Cup-shaft failure prevented calibration at higher wind velocities. The action of the wind vane was checked and found to have very poor directional accuracy below a velocity of 8 miles per hour. After shaft failure was reported to the Signal Corps, the cup rotors were redesigned by strengthening the shafts for better operation at high velocities. The anemometer with the redesigned cup rotors was recalibrated, but cup-shaft failure occurred again at a wind velocity of approximately 220 miles per hour."

Report presenting measurements of the local flow angles near the fuselages of a triangular-wing wind-tunnel model and an F-86A-5 airplane in flight using airflow detectors on the fuselages. A comparison of the measured and predicted flow angles is provided. Results indicated that the accuracy of the predicted indicated angles of attack was not sufficient to eliminate the necessity of a flight calibration of a detector mounted on a fuselage.

Report presenting an investigation to determine the effect on performance of two methods of throat-area modulation of two convergent plug nozzles. Data were obtained over a range of pressure ratios. Results regarding the performance of the translatable outer-shell-type plug nozzle, performance of the iris-outer-shell-type plug nozzles, sensitivity to throat-area variation, and air-flow parameter are provided.

Report presenting experimental data from calm- and rough-water landing impacts with a dynamic model of a seaplane incorporating a flat-bottom hydro-ski mounted on a rigid strut and on a shock-absorber strut. Results regarding the fixed-ski case and translating-ski case are provided.

From Summary: "The over-all component performance characteristics of a J71 experimental three-stage turbine with 97 percent design stator areas were determined over a range of speed and pressure ratio at inlet-air conditions of approximately 35 inches of mercury absolute and 700 degrees R. The turbine break internal efficiency at design operating conditions was 0.877; the maximum efficiency of 0.886 occurred at a pressure ratio of 4.0 at 120 percent of design equivalent rotor speed. In general, the turbine yielded a wide range of efficient operation, permitting flexibility in the choice of different modes of engine operation."

"An experimental investigation of the internal-flow conditions of a J71 experimental turbine equipped with 97-percent-design stator areas was conducted at equivalent design speed and near equivalent design work. The results of the investigation indicate that the stage work distribution closely approximates design, the actual distribution being 44.1, 33.4, and 22.5 percent for the first, second, and third stages, respectively. The first-, second-, and third-stage efficiencies were 0.894, 0.858, and 0.792, respectively" (p. 1).

Report presenting an investigation using the NACA wing-flow method at transonic speeds to determine the contributions of wing, tail, and fuselage to the aerodynamic characteristics of a semispan airplane model with a long slender fuselage and straight wing and tail of low aspect ratio with faired symmetrical double-wedge airfoil sections 4.6 percent of the chord in thickness. Results regarding the drag, pitching moment, downwash, and lift are provided.

Report presenting an investigation at transonic speeds to determine the control characteristics of a linked flap and spoiler. The control consisted of a linked combination of a quarter-span inboard plug-type spoiler and a full-span flap. Results regarding the hinge-moment, pitching-moment, lift, and rolling-moment are provided.

Report presenting an investigation of an axial-flow compressor rotor with an inlet hub-tip radius ratio of basically zero. Results regarding the radial distribution of total-pressure ratio and efficiency, measured pressure ratio at design speed, weight flow, and low-speed cascade data are provided.