Abstract: An application of the magnetic mirror principle to the acceleration of a plasma is described. It is shown that an axially symmetric magnetic field which increases with time but decreases with distance along the axis can impart a net translational energy to a plasma. This effect on a plasma is contrasted with that arising from an impressed electric field, which is not effective in producing acceleration.

Report presenting lift and drag data obtained during the Douglas X-3 airplane. The data covered the Mach number range from 0.82 to 1.20 with considerable variation in lift. A comparison of the flight data with data from wind-tunnel and rocket-model tests shows that the model tests adequately predict the performance of the airplane.

"An approximate thermal equation was derived for quenching distance based on a previously proposed diffusional treatment. The quenching distance was expressed in terms of the thermal conductivity, the fuel mole fraction, the heat capacity, the rate of the rate-controlling chemical reaction, a constant that depends on the geometry of the quenching surface, and one empirical constant. The effect of pressure on quenching distance was shown to be inversely proportional to the pressure dependence of the flame reaction, with small correction necessitated by the effect of pressure on flame temperature" (p. 1).

Memorandum presenting the lift, pitching-moment, and drag characteristics of a missile configuration with a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 at a Mach number of 1.99 and a Reynolds number of 6.0 million. Testing was performed to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components. Results regarding the body-wing-tail combination, body, wing and tail, body-wing and body-tail combinations, and body-wing-tail interference are provided.

Memorandum presenting an investigation conducted to reduce the pressure drop in an experimental combustor designed to operate with high efficiencies at high altitude. The combustor utilized a previously designed prevaporizing fuel system that supplied vapor fuel to the injectors for high-altitude operation. Results regarding combustor development and performance of final combustor model 47L are provided.

Memorandum presenting an investigation of the dynamics of a turbine-propeller engine in the altitude wind tunnel employing the frequency-response technique for a range of pressure altitudes form 10,000 to 30,000 feet. The investigation showed that the dynamic responses generalized for pressure altitude over the range of frequencies investigated. Results regarding steady-state characteristics, the linearity investigation, and dynamic characteristics are provided.

The effect of spark repitition rate on the altitude ignition limits of a single tubular (turbojet engine) combustor was investigated. An increase in sparking rate from 3 to 140 sparks per second reduced the ignition limiting combustor-inlet pressure about 2 to 4 inches of mercury for air-flow rates of 1.87 and 2.80 pounds per second per square foot.520::At 3.75 pounds pe At 3.75 pounds per second per square foot, the corresponding reduction was 4 to 12 inches of mercury. The trend was similar for both low-and high-volatility fuels and for two spark-energy levels.

Report presenting testing to obtain the effects of changes in Mach number, thickness ratio, and angle of attack on the amplitude of the pressure pulsations on several airfoils at transonic speeds. The tests were performed on NACA 65A-series airfoils with thicknesses ranging from 4 to 12 percent chord at a range of Mach numbers and angles of attack. The relations of pressure pulsations to buffeting are also provided.

Memorandum presenting four wings of different plan forms equipped with plain spoiler ailerons tested at low speeds. Three of the models had wings of aspect ratio 3 and different taper ratios and sweep of the quarter-chord lines, while the fourth had an aspect ratio of 4.8, a taper ratio of 0.51, and sweep of 35 degrees.

Report presenting four wings of different plan form equipped with plain spoiler ailerons tested at low speeds. Three of the models had wings of aspect ratio 3 and a range of taper ratios and sweep. The fourth model had aspect ratio 4.8 with taper ratio 0.51 and 35 degrees of sweep.

Report presenting the lift, pitching-moment, and drag characteristics of a missile configuration with a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 at Mach number 1.99 and Reynolds number 6.0 million. Testing occurred over a range of angles of attack and to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components.

Abstract: Uranium-vanadium ore has long been known in the Trachyte district. Production started in 1913 and has been erratic and intermittent to date. Until recently, practically all production from the Henry Mountains area has come from the Trachyte district.

Report of geological investigations of the Trachyte District of the Henry Mountains in Utah

A report about anomalous low temperature thermal conductivity of grapnite samples.

A small source model for neutron flux distribution in heterogeneous reactors is proposed. Such a model takes more explicit account of fuel lumping than has been customary in most pile calculations. The small source is applied to the cases of plate and rod fuel elements. The chief results are equations which give the pile buckling in terms of the distance between fuel elements and in terms of fuel element constants. (W.D.M.)

An annular prevaporizing turbojet combustor having pressure losses lower than those obtained in current turbojet combustors was developed, Pressure losses of 2 to 4 percent, satisfactory temperature profiles, and combustion efficiencies of 98, 88, and 81 percent were obtained at 56,000, 70,000, and 80,000 feet respectively, for a simulated 5.2- pressure-ratio engine at rated speed and 0.6 flight Mach number with JP-4 fuel. Use of JP-5 fuel resulted in a small penalty in efficiency due, at least in part, to insufficient prevaporizer capacity.

Report presenting a method to provide a simple means of comparing engine components on the basis of either range or the margin of thrust over drag. The equations include the variation in thrust coefficient and specific impulse and the change in engine weight and drag. Four general cases are considered: fixed-size airplanes with constant gross weight, fixed-size airplanes with variable gross weight, variable-size airplanes with constant payload weight, and variable-size airplanes with constant ratio of payload to gross weight.

From Introduction: "It is the purpose to describe a method of predicting spoiler rolling-moment effectiveness based on the simplified lifting-surface flap theory of reference 12. The results of applying the present method to the configurations described in references 1 to 8 (see table I and fig. 1) and the comparison with the experimental data are presented herein."

Turbine centrifugal stress is a limiting factor for all flight conditions studied. This stress is more severe for sea-level operations than for subsonic flight at the tropopause. Turbines designed for a stress of 30,000 psi are capable of driving a light, compact, high-spedd compressor but only at high values of specific fuel consumption. An increase in turbine-inlet temperature is accompanied by an increase in turbine centrifugal stress.

An investigation was made at free-stream Mach numbers of 1.62, 1.93, and 2.41 to determine the effects of a primary jet and secondary air flow on the base pressure and pressures acting over the boattailsurface of a body of revolution for two secondary discharge areas. The Mach numbers of the primary nozzles were 1 and 3.23 with the secondary mass flow being varied from 0 to 10 percent of the primary mass flow. The ratio of jet stagnation temperature to tunnel stagnation temperature was about 0.96. The Reynolds number range of the investigation was from 2.1 x 10(6) to 2.9 x 10(6)based on body length. All testing was conducted with a turbulent boundary layer along the model. This report presents results obtained with zero-length ejector and covers jet static-pressure ratios from the jet-off condition to a maximum of about 128 for the sonic nozzle and to a maximum of about 9 for the supersonic nozzle.