"A study has been made of a flare-cylinder configuration to investigate its feasibility as a reentry body of an intermediate range ballistic missile. Factors considered were heating, weight, stability, and impact velocity. A series of trajectories covering the possible range of weight-drag ratios were computed for simple truncated nose shapes of varying pointedness, and hence varying weight-drag ratios" (p. 1).

Memorandum presenting an investigation of hemispherical nose shapes of titanium, stainless steel coated with aluminum oxide, and uncoated stainless steel with lithium hydride and magnesium as internal coolants. Results regarding titanium models, stainless-steel models (uncoated), stainless-steel models coated with aluminum oxide, solution effects on the decomposition temperature of lithium hydride, effect of lithium hydride and magnesium on temperature measurements, and efficiency of models cooled with lithium hydride are provided.

Report presenting testing of hemispherical nose shapes of titanium, stainless steel coated with aluminum oxide, and uncoated stainless steel with lithium hydride and magnesium as internal coolants. The models were tested in the ceramic-heated jet at a Mach number of 2, a stagnation temperature of 4,000 degrees Fahrenheit, and a stagnation pressure of 105 lb/sq. in. abs. Lithium hydride offers significant cooling for the materials while magnesium shows a slight cooling.

"The mean strain rate in turbulent shear flow must tend to make the structure anisotropic in all parts of the spectrum. It is argued here, however, that if the spectral energy transfer process destroys orientation, the Kolmogoroff notion of local isotropy can still be justified in spectral regions where the local transfer time is shorter than the characteristics time of the gross shear strain" (p. 1).

Small metal spheres of various densities were fired at high speed into thick targets of copper and lead. In general, it was found that all of the penetrations could be correlated quite well for engineering purposes by a function relating the depth of penetration to the impact momentum per unit volume.

Memorandum presenting force-oscillation tests made in the 8-foot transonic pressure tunnel to investigate some effects of an overhang-type aerodynamic balance and of a flow spoiler on the dynamic hinge-moment characteristics of a full-span flap-type rudder on a 5-percent-thick, swept vertical fin of low aspect ratio. Test results how that the aerodynamic damping moment on the plain rudder becomes unstable near a Mach number of 0.975 and remains unstable to the maximum speed of the tests.

Report presenting force-oscillation testing in the transonic pressure tunnel to investigate some effects of an overhang-type aerodynamic balance and of a flow spoiler on the dynamic hinge-moment characteristics of a full-span flap-type rudder on a 5-percent-thick, swept vertical fin of low aspect ratio. Tests were performed at a variety of Mach numbers, reduced frequencies, and tunnel stagnation pressures, but at a constant oscillating amplitude. Results regarding aerodynamic damping, aerodynamic spring, control effectiveness, and effect of Reynolds number are provided.

Memorandum presenting an investigation conducted to measure and compare the total-pressure recovery and distortion characteristics at the compressor face of two single-place fighter-type airplanes with similar two-spool turbojet engines, but with dissimilar inlets. The total-pressure recovery was relatively independent of angle of attack and mass-flow ratio for both airplanes except for a significant decrease in pressure recovery with angle of attack for airplane B at the highest Mach numbers tested.

Report presenting an investigation to measure and to compare the total-pressure recovery and distortion characteristics at the compressor face of two single-place fighter-type airplanes with similar two-spool turbo-jet engines, but with dissimilar inlets. Results regarding compressor-face total-pressure surveys and surges encountered are provided.

Report presenting an investigation to determine the longitudinal characteristics of an airplane model with a thin, highly swept and tapered wing of low aspect ratio equipped with plain leading-edge flaps in conjunction with blowing-type boundary-layer control applied to flap radius. Several leading-edge configurations and boundary-layer control system variables were also investigated.

Report presenting an investigation to determine some effects of surface roughness on heat transfer at the stagnation point. Testing occurred in the ceramic-heated jet laboratory model at a Mach number of 2, a stagnation temperature of 3,530 degrees Fahrenheit, and a stream Reynolds number of 2.5 x 10(exp 6) per foot. Results regarding the variation of stagnation-point heat-transfer coefficient with surface roughness, heat transfer at the stagnation point of a hemisphere, and heat-transfer coefficient are provided.

Memorandum presenting a theoretical analysis of the supersonic flow about two-dimensional and three-dimensional axially symmetric models restricted by theoretical walls capable of removing the nose shock. Results regarding the supersonic-tunnel interference due to nonreflecting walls and supersonic-tunnel interference due to porous walls are provided.

Report presenting a theoretical analysis of the supersonic flow about two-dimensional and three-dimensional axially symmetric models restricted by theoretical walls capable of removing the nose shock. Results regarding the supersonic-tunnel interference due to nonreflecting walls and supersonic-tunnel interference due to porous walls are provided.

Memorandum presenting an investigation in the 4- by 4-foot supersonic pressure tunnel to determine the effects of forebody strakes on the aerodynamic characteristics in sideslip of a delta-wing airplane model at Mach number of 2.01. The presence of the strakes increased the directional-stability level for both vertical-tail arrangements. Results of pressure tunnels for a forebody show that the presence of the strakes provides a stabilizing influence on the forebody which is consistent with the results of force tests.

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).

Report presenting an investigation in the 10- by 10-foot supersonic wind tunnel to evaluate the performance of a shrouded injector burner with perforated domes employed in a 28-inch ramjet using gaseous hydrogen as fuel. Steady-state data were obtained at a pressure altitude of 77,000 feet and zero angle of attack. Results indicated that burning could be initiated under severe distortion conditions and that satisfactory combustor operation was accomplished up to a pressure altitude of 110,000 feet with no adverse effect on combustion efficiency.

Memorandum presenting an investigation conducted in the 10- by 10-foot supersonic wind tunnel to evaluate the performance of a shrouded injector burner with perforated domes employed in a 28-inch ramjet using gaseous hydrogen as fuel. Steady-state data were obtained at a pressure altitude of 77,000 feet and zero angle of attack. Results of the investigation showed that burning could be initiated under severe distortion conditions and that satisfactory combustor operation was accomplished up to a pressure altitude of 110,000 feet with no adverse effect on combustion efficiency.

Report presenting an investigation at Mach numbers of 1.61 and 2.01 to determine the pressure distributions for a series of 20 controls on a 60 degree delta wing. Tests occurred at a range of angles of attack and control deflections. Results regarding basic pressure distributions, comparison of experimental and theoretical results, and experimental comparisons are provided.

Report presenting an investigation at two Mach numbers to determine the pressure distributions for a series of 20 controls on a 60 degree delta wing. Thirteen controls were of the balanced tip type and seven of the controls were of the more conventional trailing-edge type. Results regarding the basic pressure distributions, comparison of experimental and theoretical results, and some experimental comparisons are provided.