Report discussing the use of hot-wire anemometers to measure circumferential traces of specific-mass-flow variation for various radial positions at the rotor exit. Results regarding the rotor blade-wake traces, relative total-pressure ratio, and centrifuging of low-momentum fluid are provided.

Report presenting preliminary tests of a counterrotating axial-flow supersonic compressor, which indicates that a flow limitation in the second rotor restricted the flow rate at design speed. The second-stage rotor was modified by increasing the flow area to increase the flow rate and improve flow stability. Results regarding the overall performance of the modified compressor, radial variation of flow conditions and performance parameters in the second rotor, stability of the shock wave in the second rotor, and individual stage performance are provided.

Report presenting performance of a two-dimensional external-compression inlet designed for various methods of inlet-engine matching up to Mach 3 at Mach 1.89 and Mach 0. Results regarding shock geometry, two-shock ramps, isentropic ramps, profiles, and cowl drag for the various Mach number tests are provided.

Report presenting the performance of a two-dimensional external-compression inlet designed for various methods of inlet-engine matching up to Mach 3 at Mach 1.89 and Mach 0.

Report presenting the development of a method based on blade-element concepts to evaluate the effect of stall or choke-free angle-of-attack range on the attainable pressure rise and relative Mach numbers for cascades or single blade rows. The method was extended to evaluate the effect of usable compressor-stage flow range on the design parameters of the stage flow per unit area, pressure ratio, and rotative speed.

"Theoretical rocket performance for frozen composition during expansion was calculated for liquid methane with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. The maximum calculated value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827atm) and an exit pressure of 1 atmosphere is 315.3 for 79.67 percent fluorine in the oxidant" (p. 1).