Report presenting an investigation at low speed to determine the static aerodynamic forces and moments on a canard missile model during launching from the midsemispan and wing-tip locations of a wing-fuselage combination with a 45 degree sweptback wing. When the missile was mounted under the wing at the midsemispan location, changes in chordwise position produced large changes in missile forces and moments.

Report presenting an investigation to determine the effect of blade-section camber on the stall-flutter characteristics of three propellers using a 6000-horsepower propeller dynamometer. Up to a certain blade angle, there was an increase in flutter-speed coefficient with an increase in blade-section design lift coefficient. An increase in thrust was found to lead to an increase in tip Mach number in the flutter region of the propeller blades.

Data obtained from the first flight test of a ram jet utilizing a magnesium slurry fuel are presented. The ram jet accelerated from a Mach number of 1.75 to a Mach number of 3.48 in 15.5 seconds. During this period a maximum values of air specific impulse and gross thrust coefficient were calculated to be 151 seconds and 0.658, respectively. The rocket gas generator used as a fuel-pumping system operated successfully.

Report discussing the results of an investigation to determine the stability and control characteristics of a flying model of an unswept-wing vertically rising airplane. Information about stabilization, maneuverability, rolling motions, and landings is provided.

Report presenting lift and drag data obtained during the Douglas X-3 airplane. The data covered the Mach number range from 0.82 to 1.20 with considerable variation in lift. A comparison of the flight data with data from wind-tunnel and rocket-model tests shows that the model tests adequately predict the performance of the airplane.

Report presenting the effect of channel geometry on flame quenching, as calculated on the basis of average active particle chain lengths, is related among six different geometries: plane parallel plates of infinite extent, cylindrical tubes, rectangular slots, cylindrical annuli, and tubes of elliptical and equilaterally triangular shape. The results indicated that the observed variation of flame quenching as a function of quenching geometry may be successfully predicted for a range of pressures and for rich as well as lean propane-air flames.

An experimental investigation of the performance of a single-conical-shock diffuser was conducted at a Mach number of 5.4 and a Reynolds number based on model diameter of 375,000. Total-pressure recoveries of 13.7 and 13.1 percent were obtained at angles of attack of 0 degrees and 3 degrees, respectively. The corresponding kinetic energy efficiencies were 86.4 percent at an angle of attack of 0 degrees and 86.0 percent at an angle of attack of 3 degrees.

Memorandum presenting an investigation of the performance of inlet configurations with a forebody designed for isentropic external compression at a nominal Mach number of 5.6. At zero angle of attack, all the configurations yielded larger total-pressure recoveries than had previously been obtained with a single-conical-shock inlet. Results regarding the flow about the forebody, inlet performance, and performance comparisons are provided.

Memorandum presenting an investigation of the performance of inlet configurations with a forebody designed for isentropic external compression at a nominal Mach number of 5.6. At zero angle of attack, all of the configurations yielded larger total-pressure recoveries than had previously been obtianed with a single-conical-shock inlet. Results regarding the flow about the forebody and performance comparisons are provided.

Report presenting tests in the 11-inch hypersonic tunnel at Mach number 6.86 to determine the effectiveness of an adjustable supersonic diffuser without boundary-layer control in reducing the pressure ratios across the system required to maintain supersonic flow. Testing indicated that the pressure ratio could be reduced to about one-third the pressure ratio required for starting.

Report presenting an investigation to determine the effectiveness of studying the dynamic characteristics of a low-speed slowly spinning fin-stabilized rocket with a free-oscillation technique and to study certain types of behavior that have been observed in this type of missile. Results regarding the stability of the original model, the effect of reversing direction of arming propeller, the effect of increasing the size of the reversed-rotation arming propeller, the effect of removing the arming propeller, the effect of adding a spoiler nose ring to a basic model, and the effect of changes in the spin rate are provided.

Report presenting an investigation made in the stability tunnel to determine the effectiveness with which the dynamic characteristics of a low-speed slowly spinning fin-stabilized rocket could be studied by a free-oscillation technique and to study certain peculiarities of behavior. Results regarding the stability of the original model, effect of reversing direction of the arming propeller, effect of increasing size of reversed-rotation of the arming propeller, effect of removing the arming propeller, effect of adding a spoiler nose ring to the basic model, and the effect of changes in spin rate are provided.

Report presenting a scale rocket-propelled model of the Convair XF2Y-1 water-based fighter airplane in free flight in a range of Mach numbers. Results regarding stability characteristics, trim characteristics, drag characteristics, and inlet and duct characteristics are provided.

Report presenting an investigation of two versions of a type of forward underslung scoop mounted on a pointed fuselage in the supersonic blowdown tunnel. The scoop entrances were in the shape of a 60 degree sector of an annulus, but the sides of one scoop were sweptback, while the others were unswept. Results regarding the shadowgraphs and surface-pressure distributions and total-pressure recovery are provided.

Report presenting an investigation in the transonic tunnel on a submerged inlet with a divergent-walled approach ramp to determine flow phenomena, pressure recovery, and external forces to provide correlation with available data obtained on similar inlets. Results regarding the pressure characteristics, aerodynamic forces, and performance comparisons are provided.

Report presenting a transonic investigation of the static-pressure fluctuations in the left duct of a scale inlet model of a supersonic fighter-bomber airplane in the transonic tunnel. Results regarding amplitude characteristics of pressure fluctuations, power-spectral analyses, and flow distortions at compressor face are provided.

Report presenting a transonic investigation of the static-pressure fluctuations in the left duct of a scale inlet model of a supersonic fighter-bomber airplane. The amplitude and frequency of pressure fluctuations were measured using an electrical pressure gage mounted on the diffuser wall at several Mach numbers and mass-flow ratios. Results regarding amplitude characteristics of pressure fluctuations, power-spectral analyses, and flow distortions at compressor face are provided.

Report presenting testing of wing-body combinations incorporating several wing plan forms of theoretical interest at Mach numbers of 1.62, 1.93, and 2.41 in the supersonic tunnel. One triangular, one arrow, and tow diamond plan forms were tested. Results regarding minimum drag, lift-curve slope, center-of-pressure positions, drag due to lift, and lift-drag ratios are provided.

Report presenting a determination of the performance and flow distributions in the first, fifth, and tenth stages of a 10-stage axial-flow research compressor at design and part speed in order to study the validity of the compressor design assumptions and to show the effects of part-speed compressor operation on stage performance.

Report presenting an investigation in the 20-foot free-spinning tunnel to determine the motions and trim conditions from 0 to 90 degrees angle of attack for the X-3 airplane, which is a supersonic airplane with an extremely long nose and s mall wing. Testing was conducted on several scales of models and component parts. Results from dynamic and static tests are provided.

Report discussing testing in the low-turbulence tunnel to determine the effectiveness of several configurations of vortex generators in increasing maximum lift of a 6-percent-thick symmetrical circular-arc airfoil section. None of the configurations tested were found to greatly increase the lift. Some of the configurations tested included a single row of vortex generators, two rows of generators, larger vortex generators, and sweeping forward the front row of generators.

A typical inlet axial-flow compressor inlet stage, which was designed on the basis of constant total enthalpy with symmetrical velocity diagram at all radii, was investigated. At a tip speed of 1126 feet per second, a peak pressure ratio of 1.28 was obtained at an efficiency of 0.76. At a tip speed, the highest practical flow was 28 pounds per second per square foot frontal area with an efficiency of 0.78. Data for a rotor relative inlet Mach number range of from 0.5 to 0.875 indicates that the critical value for any stage radial element is approximately 0.80 for the stage investigated.

Report presenting the results of a series of qualitative studies with tufts on three wings with sweepback angles of 30 degrees, 45 degrees, and 60 degrees, which show that under certain conditions, a leading-edge vortex can exist in the unsteady flow associated with a gust. The results indicated that if a wing in steady flight prior to entering a gust is at an angle of attack several degrees less than that at which vortex flow first begins in steady flow, it may penetrate the gust without having the vortex develop, even if its angle of attack is increased by the gust into the vortex-flow regime.

Report presenting an analysis of a series of wing-body combinations in the 8-foot high-speed tunnel. The paper presents the results of an investigation of a wing-fuselage combination with a wing of unswept quarter-chord line, aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil section. Results regarding force and moment characteristics and wake and downwash characteristics are provided.