"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).

Performance and boundary-layer data were taken in a 12 degree 10-inch inlet-diameter conical diffuser of 2:1 exit- to inlet-area ratio. These data were taken for two inlet-boundary-layer conditions. The first condition was that of a thinner inlet boundary later (boundary-layer displacement thickness, delta* approximately equal to 0.034) produced by an inlet section approximately 1 inlet diameter in length between the entrance bell and the diffuser. The second condition was a thicker inlet boundary layer (delta* approximately equal to 0.120) produced by an additional inlet section length of approximately 6 diameters. Longitudinal static-pressure distributions were measured fro wall static orifices. Transverse total- and static-pressure surveys were made at the inlet and exit stations. Boundary-layer velocity distributions were measured at seven stations between the inlet and exit. These data were obtained for a Reynolds number (based on inlet diameter) range of 1 x 10(exp 6) to 3.9 x 10(exp 6). The corresponding Mach number range was from M = 0.2 to choking. At the maximum-power-available condition supersonic flow was obtained as far as 4.5 inches downstream from the diffuser inlet with a maximum Mach number of M approximately equal to 1.5. The total-pressure loss through the diffuser in percentage of inlet dynamic …

Report presenting flight measurements of aerodynamic tail loads made on the Douglas D-558-II from which the variation with Mach number of the wing-fuselage aerodynamic center, static-longitudinal-stability parameter, the tail load per g, and the zero-lift wing-fuselage pitching-moment coefficient determined up to Mach number 0.87.

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."

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 equation for average heat transfer of a surface was derived when the boundary layer changed from laminar to turbulent. Influences on the heat transfer through a laminar boundary layer of Mach number, temperature ratio (gas temperature divided by wall temperature), and exponents of gas-property temperature relations were shown to be relatively small for air with Mach numbers less than 2 and temperature ratios between 1 and 4. Good agreement was obtained with experimental results from cylinders, an airfoil, and turbine blades.

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."

The effect of fuel volatility on altitude starting limits of an axial-flow-compressor-type turbojet engine was investigated using fuels with Reid vapor pressures of 1.1 and 5.4 pounds per square inch. At flight Mach numbers from 0.40 to 0.85, the AN-F-58 fuel allowed consistent windmilling at altitudes 2000 to 8000 feet higher than was obtained with the 1.1-pound Reid vapor pressure fuel. At a flight Mach number of 0.25, ignition could not be established at any altitude with the lower-volatility fuel.

Report presenting results from a preliminary flight investigation on a swept-wing airplane of the lateral-control characteristics and wing-dropping tendency encountered at high Mach numbers. Measurements of aileron and rudder-control effectiveness are presented and used with estimated damping-in-roll characteristics and data from steady sideslips to approximate the variation of effective dihedral with Mach number.

Summary: The Regnault Method as applied to tritium purity determinations is described. The operating difficulties, such as temperature and humidity changes, static charges, excessive vibrations, contamination with air and helium, etc., and the methods currently in use for minimizing such difficulties, are discussed. A discussion of ion-chamber calibrations, and the correlation of ion-chamber results with Regnault Method results, are included. A precision of ±1% has been obtained, but the accuracy is not yet known.

"An analytic method for the design of automatic controls is developed that starts from certain arbitrary criterions on the behavior of the controlled system and gives those physically realizable equations that the control system can follow in order to realize this behavior. The criterions used are developed in the form of certain time integrals. General results are shown for systems of second order and of any number of degrees of freedom. Detailed examples for several cases in the control of a turbojet engine are presented" (p. 207).

The purpose of this note is to recalculate the semi-empirical curve of neutron yield vs. deuteron energy originally obtained by H. York in order to predict, using it and the best available data, what yields may be expected for deuterons on uranium in the extrapolated energy region.

Report presenting an investigation of the design characteristics and loadings conducive to the tumbling of airplanes with no horizontal tails. Empirical design requirements based on model testings of 18 different configurations and an explanation of the phenomena of tumbling are provided.

Report presenting results of a study regarding the critical masses and nuclear properties of plutonium as well as the use and effects of various tampers.

Memorandum presenting flight tests made on the Bell X-1 airplane with the 10-percent-thick wing and the 8-percent-thick tail to evaluate the effects of fixing the rudder and changing the inclination of the principal axes of inertia. The data were obtained during power-off glides at altitudes between 32,000 and 16,000 feet.

Report presenting flight measurements of the dynamic longitudinal-stability characteristics of a 35 degree swept-wing airplane. The results indicated that a sharp reduction in damping at a Mach number of 0.92 with maximum damping occurred at Mach number of 0.88 and 0.94. Results regarding oscillation periods, damping coefficients, scatter, and number of cycles required for oscillations to damp to a certain amplitude are provided.

Report presenting an investigation of a model of the Bumblebee XPM missile to determine the causes of booster-fin failures for a range of Mach numbers and Reynolds numbers. The cause of fin failure was determined to be the launching shoes that caused the missiles to trim at increasingly negative angles of attack. Additional testing with wing spoilers and alternate booster fins was also conducted.

Report presenting the results of an investigation of a model of the Bumblebee XPM missile to determine the causes of booster-fin failures on full-scale missiles. Results indicated that failures of the fins were due to launching shoes which caused the missile to try at increasingly negative angles of attack as the Mach number increased.

"Two theoretical procedures are developed for designing asymmetric supersonic nozzles for which the calculated exit flow is nearly uniform over a range of Mach numbers. One procedure is applicable at Mach numbers less than approximately 3. This approach yields, without iteration, a nozzle for which the calculated exit flow is uniform at two Mach numbers and, with proper design, is nearly uniform at Mach numbers between, slightly above, and slightly below these two" (p. 1).

Report discussing testing to determine the sensitivity to sideslip of the static-pressure orifices of the Boeing B-29. The purpose was to determine the errors in indicated altitude and vertical velocity that occur when an airplane experiences oscillations in sideslip on a bombing run. The sensitivity of the orifices to sideslip was determined by measurements of pressure and sideslip angle during constant-amplitude oscillations in sideslip.

From Introduction: "The primary purpose of the investigation described in the present described in the present report is to formulate a method which is of value for qualitative calculations of base pressure both on airfoils and bodies."

The state of polonium in aqueous solutions was investigated by means of ion exchange equilibrium studies. The distribution of the polonium between the solid and liquid phases was determined for solutions of various compositions using both a cation and an anion exchanges. These studies show that in the absence of a complexing agent polonium is hydrolyzed to such a degree that it forms monovalent cations and anions whose charge was not investigated in this study.

One of the problems connected with the sweptback wing is the difficulty of controlling the location of the center of pressure and hence the pitching moment. A method is presented for designing a wing to be self-trimming at a given set of flight conditions. Concurrently, the spanwise distribution of load on the wing is made to be approximately elliptical, in an effort to maintain low wing drag.

"An experimental investigation of the variation of aileron rolling effectiveness and total drag with Mach number has been made using 1/6-scale rocket-propelled models of the Bell MX-776. Three models having constant-chordwise-thickness full-span aileron at approximate deflections of 2 deg, 5 deg, and 15 deg have been flown. Positive control effectiveness over the Mach number range between approximately 0.5 and 1.2 was obtained from the models and no indication of reversal of effectiveness was encountered" (p. 1).