Memorandum presenting an investigation to determine the drag, static longitudinal and lateral stability, and longitudinal trim characteristics of an airplane configuration with tail surfaces outboard of the wing tips. Included in the basic data are some effects of Reynolds number, engine pack, and wing twist combined with toe-out of the vertical tails.

Memorandum presenting a two-dimensional wind-tunnel investigation to determine the extent to which boundary-layer removal through slots is effective as a means for maintaining extensive laminar layers at high Reynolds numbers. Results regarding drag, suction-flow and pressure-loss distribution, surface pressure distribution, lift, and difficulties encountered in obtaining extensive laminar flow are provided.

Report presenting results obtained to determine the effects of compressibility at high Mach numbers on a scale model of the Bell XS-1 transonic research airplane. The results are presented for several model configurations through a range of Mach numbers. Results regarding aerodynamic characteristics and static longitudinal stability characteristics are provided.

Report presenting an investigation to determine the effect of controlled grain size and increased carbon content on engine performance and high-temperature properties of precision-cast Stellite 21 turbine blades. Results regarding blade stresses, engine operation, and metallurgical examination are provided.

Report presenting a flight investigation to determine the zero-lift drag characteristics of a twin-engine supersonic ramjet configuration in order to make performance calculations below the design Mach number. Data was obtained over a range of Mach numbers for two power-off ramjet models with and without nacelles. Information about the variation of Reynolds number, total and external drag coefficients, and model-component drag coefficients is provided.

Investigation of the internal performance characteristics of three types of variable-throat convergent-divergent nozzles over a wide range of pressure ratios. Only small differences in peak thrust coefficient between the three types of variable-throat convergent-divergent nozzles were obtained.

Report presenting an investigation of several two-dimensional supersonic inlet configurations at a Mach number of 2.01 and an angle of attack of 0 degrees. Two inlets similar in supersonic diffuser design but one with a subsonic diffusers about half the length of the other one were also investigated. Results regarding the inlet design, pressure recovery, shadowgraphs, and suggestions for inlet improvement are provided.

Report presenting an investigation to determine the normal force of two three-blade NACA propellers and the agreement of the measured normal force with a theoretical method. Propeller thrust and normal force were measured for a variety of blade angles, angles of attack, and Mach numbers. Results regarding variation with thrust, effect of inclination, and a comparison with theory are provided.

"An investigation at a free-stream Mach number of 2.02 was made to determine the effects of a propulsive jet on a wing surface located in the vicinity of a choked convergent nozzle. Static-pressure surveys were made on a flat surface that was located in the vicinity of the propulsive jet. The nozzle was operated over a range of exit pressure ratios at different fixed vertical distances from the flat surface. Within the scope of this investigation, it was found that shock waves, formed in the external flow because of the presence of the propulsive jet, impinged on the flat surface and greatly altered the pressure distribution" (p. 1).

"Extensive experimental measurements have been made of the downwash angles and the wake characteristics behind airfoils with and without flaps and the data have been analyzed and correlated with the theory. A detailed study was made of the errors involved in applying lifting-line theory, such as the effects of a finite wing chord, the rolling-up of the trailing vortex sheet, and the wake. The downwash angles, as computed from the theoretical span load distribution by means of the Biot-Savart equation, were found to be in satisfactory agreement with the experimental results" (p. 179).

Report presenting the flutter of a solid wing as affected by aerodynamic heating, which can cause a large momentary loss in torsional stiffness. Both experimental and analytical studies were conducted and good correlation was demonstrated. The loss of stiffness due to aerodynamic heating was calculated and the operational line intersected with the flutter curve.

"A survey has been made of some of the factors to be considered in the design of canard aircraft configurations. The factors include Mach number and angle-of-attack effects as well as the effects of various geometric changes. Among the geometric variables included are the effects of wing plan form, wing height, wing twist, canard plan form, canard area, canard moment arm, forebody length, afterbody length, forebody deflection, vertical-tail plan form, vertical-tail size, vertical-tail location, and various ventral-fin arrangements" (p. 1).

"A small-scale transonic investigation of two semispan wings of the same plan form was made in the Langley high-speed 7- by 10-foot tunnel through a Mach number range of 0.70 to 1.10 and a mean-test Reynolds number range of 745,000 to 845,000 to determine the effects of partial-span leading-edge camber on the aerodynamic characteristics of a swept-back wing. This paper presents the results of the investigation of wing-alone and wing-fuselage configurations of the two wings; one, was an uncambered wing and the other had the forward 45 percent of the chord cambered over the outboard 55 percent of the span. The semispan wings had 50deg 38ft sweepback of their quarter-chord lines, aspect ratio of 2.98, taper ratio of 0.45, and modified NACA 64A-series airfoil sections tapered in thickness ratio" (p. 1).

Memorandum presenting limited flow-field surveys made behind a large-scale thin straight wing of aspect ratio 3, both alone and in combination with a body and vertical tail. The surveys indicated a region of high downwash angles, low dynamic pressures, and rough flow, that extended higher above the wing chord plane, and farther aft, with increasing angle of attack beyond the tail.

"A heat-transfer investigation was conducted with air flowing through an electrically heated silicon carbide tube with a rounded entrance, an inside diameter of 3/4 inch, and an effective heat-transfer length of 12 inches over a range of Reynolds numbers up to 300,000 and a range of average inside-tube-wall temperature up to 2500 R. The highest corresponding local outside-tube-wall temperature was 3010 R. Correlation of the heat-transfer data using the conventional Nueselt relation wherein physical properties of the fluid were evaluated at average bulk temperature resulted in a separation of data with tube-wall-temperature level" (p. 1).

"The probable spin and recovery characteristics of the XSN2J-1 air-plane have been estimated on the basis of the results of brief test a performed on a model of an airplane of somewhat similar design. The spin-recovery tail-parachute requirements for the airplane were also determined end, in addition, an analysis was made to determine the best method of emergency pilot escape during a spin. The results of the investigation indicate that the recovery characteristics of the airplane will be satisfactory for all probable loading conditions of the airplane" (p. 1).

Report presenting a wind-tunnel investigation to determine the extent to which boundary-layer removal through slots is effective as a means for maintaining extensive laminar layers at high Reynolds numbers. Results regarding drag, suction-flow and pressure-loss distribution, surface pressure distribution, lift, and some of the difficulties involved in obtaining extensive laminar flow are provided.

A heat-transfer investigation was conducted with air flowing through an electrically heated silicon carbide tube with a rounded entrance, an inside diameter of 3/4 inch, and effective heat-transfer length of 12 inches over a range of Reynolds numbers up to 300,000 and a range of average inside-tube-wall temperatures up to 2500 R. The highest corresponding local outside-tube-wall temperature was 3010 R.

Report presenting an experimental investigation using a 1 7/8-inch diameter burner to determine the suitability of residual fuel oils as carriers in magnesium slurries. Four residual fuel oils with varying viscosities were investigated by themselves and in slurries of 50 percent magnesium. Results regarding the settling characteristics, blow-out velocity, and combustion efficiency are provided.

Report presenting testing of a 5-percent-thick 2024-T3 aluminum-alloy multiweb wing (MW-1-(2)) at a Mach number of 2 under simulated supersonic flight conditions. The wing experienced dynamic failure as a result of catastrophic flag-waving flutter due to the combined action of aerodynamic heating and loading. Results regarding model temperatures, experimental pressures, strain results, accelerometer, and model failure are provided.

"An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the static stability and control and drag characteristics of a 1/15-scale model of the Grunman F9F-9 airplane. The effects of alternate fuselage shapes, wing camber, wing fences, and fuselage dive brakes on the aerodynamic characteristics were also investigated. These tests were made at a Reynolds number of 1.96 x l0 (exp 6) based on the wing mean aerodynamic chord of 0.545 foot" (p. 1).