"An experimental investigation is being conducted at the Lewis laboratory to establish general principles for the design of noncritical turbine rotor configurations. This investigation includes evaluation of cooling effectiveness, structural stability, cooling-air flow distribution characteristics, and methods of supplying cooling air to the turbine rotor blades. Prior to design of a noncritical rotor, a standard turbine rotor of a commerical turbojet engine was split in the plane of rotation and machined to provide a passage for distributing cooling air to the base of each blade" (p. 1).

Report presenting the lift, drag, and pitching-moment characteristics of a wing-fuselage combination employing a wing with the leading edge swept back 63 degrees and cambered and twisted for a uniform load at a lift coefficient of 0.25 and Mach number 1.53. The experimental results are compared to those from other investigations. A maximum lift-drag ratio of 8.9 and a minimum drag coefficient of 0.0145 was obtained.

"An analysis is presented of the influence of wing aspect ratio and tail location on the effects of compressibility upon static longitudinal stability. The investigation showed that the use of reduced wing aspect ratios or short tail lengths leads to serious reductions in high-speed stability and the possibility of high-speed instability" (p. 1).

Report presenting a small-scale transonic investigation of a semispan wing sweptback 45 degrees and of aspect ratio 4 with combinations of nose-flap deflections and wing twist in the high-speed 7- by 10-foot tunnel over a range of Mach numbers. Lift, drag, and pitching-moment data were obtained for several different configurations.

Report presenting an investigation at subsonic speeds to determine the effects of lip shape and entrance Mach number on the characteristics of a scoop-type air-induction system designed for an airplane flying at supersonic speeds. Results regarding the ram-recovery ratio, static-pressure distribution, and drag are provided.

"Ignition delay determinations of several fuels with nitric oxidants were made at simulated altitude conditions utilizing a small-scale rocket engine of approximately 50 pounds thrust. Included in the fuels were aniline, hydrazine hydrate, furfuryl alcohol, furfuryl mercaptan, turpentine, and mixtures of triethylamine with mixed xylidines and diallyaniline. Red fuming, white fuming, and anhydrous nitric acids were used with and without additives. A diallylaniline - triethylamine mixture and a red fuming nitric acid analyzing 3.5 percent water and 16 percent NO2 by weight was found to have a wide temperature-pressure ignition range, yielding average delays from 13 milliseconds at 110 degrees F to 55 milliseconds at -95 degrees F regardless of the initial ambient pressure that ranged from sea-level pressure altitude of 94,000 feet" (p. 1).

Report presenting flutter characteristics at transonic speeds of a model of a fighter-type-airplane wing over a range of Mach numbers. The results indicated an abrupt increase in flutter speed ratio near a Mach number of 1.0, with only slight increases noted as it was increased to 1.4. Results regarding the effect of flutter-boundary shape are provided.

Characteristic velocities of six single-element injectors that separately varied propellant spreading and mixing were measured over a range of oxidant-fuel weight ratios of 2 to 7. Hydrogen inlet temperature was -320 degrees F. Injectors that both mixed and spread the propellants had the highest characteristic-velocity efficiency. Spreading of either propellant produced about the same increment efficiency. The increase in efficiency due to mixing was relatively small. With similar injectors, oxygen-hydrogen required about 0.2 to 0.5 of the combustor volume of oxygen-heptane for the same efficiency.

Report presenting eight thin magnesium fins, seven with the leading edges swept back 17 degrees and one with the leading edge swept back 45 degrees, in the preflight high-temperature jet. The investigation was made to determine the effectiveness of various protective coverings designed to alleviate aerodynamic-heating effects and intended for application on the first stage of rocket-propelled multistage hypersonic models.

"An investigation of tantalum was made in still air and in high-velocity air to determine its resistance to oxidation at high temperatures. Experiments show that, except for a narrow range of temperature, unprotected tantalum is not suitable for airframe parts exposed to air at high stagnation temperatures, because above this temperature range tantalum oxidizes or burns rapidly" (p. 1).

Report presenting a laboratory-scale vapor-deposition coating facility used in the field of high-temperature coatings. The vapor-deposition coatings were produced by the hydrogen reduction of halides and not by vacuum plating methods. The coated model was found to reach equilibrium temperature of about 3100 degrees Fahrenheit and was undamaged after 470 seconds at equilibrium temperature, while an uncoated model was destroyed in less than 6 seconds.

"An experimental investigation of combustor total-pressure loss was undertaken to confirm previous theoretical analyses of effects of geometric and flow variables and of heat addition. The results indicate that a reasonable estimate of cold-flow total-pressure-loss coefficient may be obtained from the theoretical analyses. Calculated total-pressure loss due to heat addition agreed with experimental data only when there was no flame ejection from the liner at the upstream air-entry holes" (p. 1).

Report presenting testing of three uninstrumented tapered magnesium fins with the leading edges swept back 17 degrees in an ethylene-heated high-temperature jet. The testing was carried out to investigate some effects of leading-edge diameter and leading-edge shape on the aerodynamic heating by noting the time for melting to begin on the fins. Results indicated that increasing the diameter of cylindrical leading edge increased the time required for melting to start.

Memorandum presenting testing of several materials including graphite, silicon carbide, and a number of polymer-glass-cloth laminated constructions at temperatures of 3,000 and 4,000 degrees Fahrenheit in a laboratory-scale ceramic-heated air jet. The tests were made to utilize four possible mechanisms for the alleviation of the aerodynamic-heating problem of hypersonic aircraft, which included radiative heat transfer, pyrolysis, fusion, and mass-transfer cooling by ablation.

"This report presents the results of wind-tunnel tests of a 0.22-scale model of the North American XP-82 airplane with several modifications designed to reduce the buffeting of the airplane. The effects of various modifications on the air flow over the model are shown by means of photographs of tufts. The drag, lift, and pitching-moment coefficients of the model with several of the modifications are shown" (p. 1).

Report presenting a wind-tunnel investigation of an external store of fineness ratio 8.6 on the chord plane of several low-aspect-ratio wings. Testing occurred at a Mach number of 1.9 on wings with 60 degree leading-edge sweepback and unswept trailing edges with taper ratios of 0 and 0.28 and on an unswept wing with a taper ratio of 0.625. Results regarding the loading breakdown, store position, and effect of store on control characteristics are provided.

Report presenting an investigation of a transonic turbine with a 0.6 hub-tip radius ratio in order to determine the effect of stator and rotor secondary flows on turbine performance. Results regarding the overall performance, stator exit surveys, and rotor exit surveys are provided.

A combination sliding-ring and pressure-balancing seal capable of transferring pressurize air from stationary to rotating parts was developed and experimentally investigated at sliding velocities and cooling-air pressures up to 10,000 feet per minute and 38.3 pounds per square inch absolute, respectively. Leakage of cooling air was completely eliminated with an expenditure of balance air less than one-fourth the leakage loss of air from labyrinth seals under the same conditions. Additional cooling of the carbon-base seal rings was required, and the maximum wear rate on the rings was about 0.0005 inch per hour.

An investigation has been made in the Langley 9- by 12-inch supersonic blowdown tunnel to determine the effects of external-store location on the lift, drag, and pitching-moment characteristics of a 45 degree sweptback wing at Mach numbers of 1.41, 1.62, and 1.96. The spanwise, chordwise, and vertical location of a Douglas-Aircraft Company, Inc., store of fineness ratio 8.58 was systematically varied over the outer 60 percent of the wing semispan. A brief investigation of strut sweep angle was also made. The test Reynolds number based on the wing mean aerodynamic chord ranged from 1.3 x 10(exp 6) to 1.5 x 10(exp 6).