Report presenting a blade-vibration and rotating-stall survey on a modified version of a production turbojet engine with a 13-stage axial-flow compressor with a design pressure ratio of 7 and an air flow of 120 pounds per second. This testing was carried out due to the problems rotating stall can cause with both experimental and production axial-flow compressors. Results regarding the rotating-stall patterns, rotor-blade vibrations, radial and axial strength of rotating stall, and maximum rotating-stall speed are provided.

Effective flame speeds of free turbulent flames were measured by photographic, ionization-gap, and photomultiplier-tube methods, and were found to have a statistical distribution attributed to the nature of the turbulent field. The effective turbulent flame speeds for the free flame were less than those previously measured for flames stabilized on nozzle burners, Bunsen burners, and bluff bodies. The statistical spread of the effective turbulent flame speeds was markedly wider in the lean and rich fuel-air-ratio regions, which might be attributed to the greater sensitivity of laminar flame speed to flame temperature in those regions. Values calculated from the turbulent free-flame-speed analysis proposed by Tucker apparently form upper limits for the statistical spread of free-flame-speed data. Hot-wire anemometer measurements of the longitudinal velocity fluctuation intensity and longitudinal correlation coefficient were made and were employed in the comparison of data and in the theoretical calculation of turbulent flame speed.

For a given wing and supersonic Mach number, the problem of shaping an adjoining fuselage so that the combination will have a low wave drag is considered. Only fuselages that can be simulated by singularities (multipoles) distributed along the body axis are studied. However, the optimum variations of such singularities are completely specified in terms of the given wing geometry. An application is made to an elliptic wing having a biconvex section, a thickness-chord ratio equal to 0.05 at the root, and an aspect ratio equal to 3. A comparison of the theoretical results with a wind-tunnel experiment is also presented.

Memorandum describing an investigation using two airplanes, an F6F-3 and F-86A, each fitted with servo equipment for varying in flight the lateral and directional stability and handling characteristics, which have been flown by test pilots to simulate the predicted dynamic behavior of six prototype airplanes. The methods of simulation and the types and ranges of variables considered are presented and the results of the individual programs are discussed.

Report presenting an investigation of the use of aerodynamic and mechanical devices for improving the response of guided missiles. An analysis indicates that the use of free-floating flaps and spring-mounted control surfaces should be able to increase maneuverability to the point that electronic automatic stabilization and gain-adjusting devices can be eliminated.

"Some insight into the range of applicability of the transonic area rule has been gained by comparison with the appropriate similarity rule of transonic flow theory and with available experimental data for a large family of rectangular wings having NACA 63AXXX profieles. In spite of the small number of geometric variables available for such a family, the range is sufficient that cases both compatible and incompatible with the area rule are included" (p. 1).

Report presenting pressure distributions and force characteristics for a body of revolution consisting of a fineness ratio of 5.75, circular-arc, ogival nose tangent to a cylindrical afterbody for a range of angles of attack. Comparison of the theoretical and experimental pressure distributions shows that for zero lift, both slender-body theory and higher-order theories yield results that are in good agreement with theory.

Effects of Reynolds number and angle of attack on the pressure distribution and normal-force characteristics of a body of revolution consisting of a fineness ratio 3 ogival nose tangent to a cylindrical afterbody 7 diameters long have been determined. The test Mach number was 1.98 and the angle-of-attack range from 0 degree to 20 degrees. The Reynolds numbers, based on body diameter, were 0.15 x 10(6) and 0.45 x 10(6). The experimental results are compared with theory.

Report presenting an investigation to determine the transition Reynolds numbers on a 10 degree cone in the 4- by 4-foot supersonic pressure tunnel at three Mach numbers and a range of Reynolds numbers. Results regarding a smooth cone and a cone with roughness are provided.

Note presenting a determination of the boundary layer behind a shock or thin expansion wave advancing into a stationary fluid. The assumption of a thin expansion wave was found to be valid for weak expansions but becomes progressively less accurate for strong expansion waves. The turbulent-boundary-layer solutions in this report represent an extension of empirical, semi-infinite flat-plate, boundary-layer data to the case where the wall is moving.

Report presenting a description of a temperature-measuring system using two sonic-flow orifices in series, which is used to measure the exhaust gas temperature of an afterburning jet engine mounted in a swinging pitot-static probe. It was found to improve the reliability and reduce the computational time required of the previously used system.

An investigation of the effectiveness of a wing equipped with a sliding flap and a leading-edge slat in deflecting a propeller slipstream downward for vertical take-off was conducted in a static-thrust facility. Results regarding the characteristics with leading-edge slat and with extended flap chord are provided.

A study to evaluate a method for predicting the downwash in a transverse plane behind a wing-body combination throughout a range of angles of attack. The wing-body combination had a ratio of maximum body diameter to wing span of 0.259, a triangular wing with an aspect ratio of 2.0, and a body of revolution with a fineness ratio of 12.5. Results regarding span loading, downwash, and vortex cores are provided.

Report presenting an investigation to determine the effects of the various components and combinations of components on the static longitudinal and lateral stability characteristics of unswept-midwing models having wings with a variety of aspect ratios for a range of angles of attack. Results regarding static longitudinal stability characteristics and static lateral stability characteristics are provided.

Report presenting an investigation of a leading-edge slat as a possible longitudinal control device for vertically rising airplanes that utilize the redirected-slipstream principle at zero forward speed in a static-thrust facility. A semispan wing model equipped with large-chord slotted flaps and two large-diameter overlapping propellers was used. Results regarding the effect of slat position, effect of propeller position, characteristics of the modified model, and a comparison of characteristics with one or two propellers are provided.

Report presenting an investigation of a model of the Republic F-105 airplane to determine the static longitudinal aerodynamic characteristics of a revised configuration. The revisions included a lengthened fuselage, a relocated canopy, a contoured fuselage afterbody, a ventral fin, and an enlarged vertical tail. Results regarding the effect of stores, effect of dive-brake flaps, effect of forebody modifications, gun blisters, and duct air bleed, and a comparison of the modified model with the original are provided.

Report discussing testing of a model of the McDonnell F3H-2N to determine its pitch-up and buffet boundaries and its longitudinal stability and control data obtainable with the pulse-tail technique. Stability was found to be less at low trim angles of attack than at high trim angles of attack up to a point. The buffet boundary was not obtainable through this testing.

Memorandum presenting an experimental investigation of various methods for the control of afterburner combustion screech at afterburner-inlet total pressures from 4000 to 6400 pounds per square foot absolute. Generally, the range of afterburner fuel-air ratios in which screech occurred and the intensity of screech did not vary appreciably in the range of pressures covered.

Note presenting a nonlinear plate theory of motion, valid for large deflections, in the light of the three-dimensional theory and other nonlinear plate theories. The nonlinear equations are solved for the case of propagation of straight-crested waves and the wave velocities are computed for various values of the parameters involved.

Theoretical calculations of specific impulse to determine the separate effects of increasing the combustion-chamber pressure and the nozzle expansion ratio on the performance of the propellants, hydrogen-fluorine, hydrogen-oxygen, ammonia-fluorine and AN-F-58 fuel - white fuming nitric acid (95 percent). The results indicate that an increase in specific impulse obtainable with an increase in combustion-chamber pressure is almost entirely caused by the increased expansion ratio through the nozzle.

Report presenting the unsteady-lift functions for a wing undergoing a sudden change in sinking speed for delta wings with three aspect ratios and rectangular and elliptical wings with three aspect ratios. Results regarding the delta-plan-form K1 function, indical spanwise loading distributions, effect of plan-form shape on the K1 function, and effect of aspect ratio on the K1 and K2 functions are provided.

Note presenting the effect of inclination of the airstream on the measured pressures of 54 total-pressure tubes for angles of attack up to 60 degrees and over a Mach number range from 0.26 to 1.62. The investigation was conducted in five wind tunnels. Results regarding effect of impact-opening size, effect of varying the shape of the internal chamber, effect of external shape, effect of slant profile, effect of venting of shielded tubes, effect of varying the probe position and the throat diameter of shielded tubes, effect of slant profile on shielded tubes, effect of the shape of the entry of shielded tubes, effect of Mach number on nonshielded tubes, and effect of Mach number on shielded tubes are provided.

Report presenting measurements of the oscillating air forces on a two-dimensional wing oscillating in pitch about the midchord that have been measured at various mean angles of attack and at two different Mach numbers. The magnitudes of normal-force and pitching-moment coefficients were found to be much higher at large angles of attack than at low angles of attack for some conditions. Results regarding the effects at low angles of attack, high angles of attack, and a description of finite-span wings oscillating in first bending mode are provided.

Report presenting the experimental results of an investigation to determine compressive strength and creep lifetime of 2024-T aluminum-alloy skin-stringer panels at room temperature and 400 degrees Fahrenheit. A method which makes use of time-dependent compressive stress-strain curves for predicting creep lifetime of panels is presented.