Report discussing a general introduction to heavy element chemistry.

An effort to develop a control on the air velocity at the door opening of a fume hood.

An investigation of the pressure distribution on the fuselage nose and the pilot canopy of a supersonic airplane model has been conducted at a Mach number of 1.90 over a wide range of angles of attack and yaw. Boundary layer separation apparently occurred from the upper surface at angles of attack above 24 degrees and from the lower surface at minus 15 degrees. No separation from the sides of the fuselage was evident at yaw angles up to 12 degrees.

The Allison model 400-C6 compressor was operated at an inlet pressure of 12 inches of mercury absolute ana ambient inlet temperature at equivalent impeller speeds of 6000, 7000, and 8500 rpm. Additional runs at an equivalent speed of 7000 rpm and ambient inlet temperature were made at inlet pressures from 7 to 22 inches of mercury absolute. The results of this investigation are compared with those of the 533-A-23 compressors. For the speeds investigated, the Allison model 400-C6 compressor had a maximum adiabatic temperature-rise efficiency of 0.768 at an equivalent speed of 7000 rpm; the corresponding equivalent weight flow was 45.0 pounds per second and the pressure ratio was 1.83. At an equivalent impeller speed of 8500 rpm, the maximum equivalent weight flow was 61.6 pounds per second and the peak pressure ratio of 2.38 occurred at an equivalent weight flow of 52.2 pounds per 1 second and an adiabatic temperature-rise efficiency of 0.714. At an equivalent speed of 7000 rpm, increasing the compressor- inlet pressure increased the maximum equivalent weight flow and the pressure ratio.

Report presenting a method for the design of supersonic-wind-tunnel nozzles that produce uniform flow over a continuously variable Mach number range without the use of flexible walls. Experimental results from a preliminary investigation of the tunnel are included, including calibration survey of flow, Mach number distributions, and pressure gradients are provided.

Report presenting an investigation in the 20-inch supersonic tunnel at Mach number 1.92 with a 3.6-inch-diameter ram jet to determine the effects of combustion on the performance of a shock diffuser with a triple-shock projecting cone. Results are presented for a range of outlet areas and fuel flows with both single- and split-fuel-injection systems. Results regarding diffuser performance with combustion, combustion with constant outlet-inlet area ratio, combustion with constant fuel flow, diffuser performance with combustion, and Mach number distribution at the diffuser outlet are provided.