"An investigation has been made in the Langley 4-by-4-foot supersonic pressure tunnel to determine the aerodynamic characteristics of a series of missile configurations having low-aspect-ratio wings at a Mach number of 2.01. The effects of wing plan form and size, length-diameter ratio, forebody and afterbody length, boattailed and flared afterbodies, and component force and moment data are presented for combined angles of attack and sideslip to about 28 degrees. No analysis of the data was made in this report" (p. 1).

Memorandum presenting the pressure recovery, mass flow, and axial force of four bodies with nose inlets measured at Mach numbers between 1.4 and 2.0 and angles of attack of 0, 3, 6, and 9 degrees. The drag coefficients of axially symmetric diffusers operating at the maximum mass-flow rates were calculated from schileren photographs of the head shock waves and frictional drag considerations. Results regarding open-nose diffusers and conical-shock diffusers are provided.

Report presenting the effect on drag of varying the vertical position of a nacelle at 40 percent semispan of a 45 degree sweptback wing and body combination as determined by transonic flight tests at zero lift. Three nacelle positions were investigated: a pylon mounted nacelle, an underslung nacelle, and a symmetrically mounted nacelle. Results regarding total drag coefficient and maximum wing-tip helix angle are provided.

"Aerodynamic data have been obtained for a 45-degree sweptback wing-body-canopy configuration at transonic speeds with the canopy placed on the body so that the cross-sectional area of the canopy approximately filled the concave portion of the basic wing-body cross-sectional-area distribution curve (design location) and with the canopy placed 0.0614 of the body length forward of the design location. Data have also been obtained for the basic wing-body combination. Placing the canopy in the rear position significantly reduced the drag of the configuration at transonic speeds, increased the lift, and did not appreciably affect the slope of the pitching-moment curve" (p. 1).

Memorandum presenting tests in the 9-inch supersonic tunnel of three bomb or store shapes without fins and with two sets of cruciform fins. Measurements of normal force, chord force, and pitching moment were made; analyses of the effects of angle of attack and Reynolds number on these qualities and also upon center-of-pressure position and fin effectiveness are presented.

Memorandum presenting an exploratory investigation conducted to determine the performance characteristics of a fuel-rich hydrogen combustor; in addition, the performance of an afterburner operating with the fuel-rich exhaust mixture was evaluated. Results regarding performance of the fuel-rich combustor, afterburner performance, and resonating combustion are provided.

Report presenting results and analysis of the longitudinal stability and control characteristics of the Douglas D-558-I airplane at a range of altitudes and Mach numbers. It was found that large and rapid changes in elevator deflection and force were required for balance at Mach numbers above 0.84 due to a sharp decrease in elevator-stabilizer effectiveness.

Flight tests were conducted at the Flight Test Station of the Pilotless Aircraft Research Division at Wallop Island, Va., to determine the longitudinal control and stability characteristics of 0.5-scale models of the Fairchild Lark pilotless aircraft with the tail in line with the wings a d with the horizontal wing flaps deflected 60 deg. The data were obtained by the use of a telemeter and by radar tracking.

Report presenting the results of force and moment tests at Mach numbers from 3.0 to 6.3 on bodies of revolution of fineness ratios from 5 to 10 and on flat-bottom bodies of fineness ratio 10 along with theoretical predictions for these bodies. Testing included eight cone and cone-cylinder models, six nose-cylinder models, and three flat-bottom bodies.

Memorandum presenting an investigation to evaluate the effect of Reynolds number and Mach number on the aerodynamic characteristics of a horizontal tail of aspect ratio 3 equipped with a plain, sealed, full-span elevator. The line joining the quarter-chord points of the airfoil sections was swept back 45 degrees and the sections perpendicular to this line were the NACA 64A010. Results regarding the effect of Reynolds number and effects of leading-edge roughness and elevator-nose seal are provided.

Report presenting an investigation of the low-speed power-on stability and control characteristics of a free-flying model of the North American X-15 airplane. Results regarding the longitudinal stability and control, lateral stability, and lateral control are provided. Longitudinal stability was low, but it was considered to be satisfactory up to an angle of attack of about 30 degrees.

Report presenting drag coefficients determined by the accelerometer method for the Bell X-1 airplane with 8-percent-thick wing and 6-percent-thick tail in power-off flight over a range of Mach numbers and lift coefficients. Results regarding the variation of computed lift and drag coefficients, variation of lift coefficient with Mach number, and effect of Mach number on lift-drag ratio are provided.

"An investigation was made to determine the static lateral stability and control characteristics of a 1/6-scale model of the Republic XF-84H airplane with the propeller operating. The model had a 40 degree swept wing of aspect ratio 3.45 and had a thin 3-blade supersonic-type propeller. Many modifications to the basic configuration were investigated in attempts to alleviate lateral and directional trim problems which appeared to be associated with propeller slipstream rotation" (p. 1).

Report presenting the results of an investigation conducted in the altitude wind tunnel to determine the performance of a 20-inch ramjet. The investigation was conducted at a range of altitudes and ram-pressure ratios using preheated 62-octane fuel. A net thrust of 8135 pounds educed to standard sea-level conditions, a net-thrust coefficient of 0.74, and an overall efficiency of 12.6 percent were attained at Mach number 1.845 at which the engine with a 5-foot combustion chamber was operated.

"Results have been obtained by means of a free-flight technique utilizing rocket propulsion which indicate that aileron-rolling-effectiveness characteristics are affected adversely by variations in airfoil section which procedure large increases in the trailing-edge angle" (p. 1).

Report presenting testing of a full-scale, semispan, horizontal tail surface for a jet-propelled fighter airplane in the 16-foot high-speed tunnel to determine the variation of the chordwise and spanwise pressure distributions with Mach number. Surface irregularities were found to cause appreciable distortion of the pressure distribution. Results regarding chordwise distributions, spanwise distribution, effect of Mach number on root bending moment, and estimated critical Mach numbers are provided.

Report presenting measurements of pressure made over the chord of a swept-wing model at four stations along the span by the NACA wing-flow method. The model was an untapered 45 degree sweptback wing of aspect ratio 3.5 with 2-inch-chord NACA 65-210 airfoil sections normal to the leading edge. Results regarding the chordwise pressure distributions and spanwise distributions of section lift, drag, and pitching moment.

Pressure distributions were measured over rectangular wings of aspect ratio 2 and triangular wings of aspect ratios 2 and 4 at Mach numbers of 1.45 and 1.97. The investigation includes some comparison of the effects of Reynolds number and of thickening the wing root sections on the loading.

"An attempt is made to describe the phenomenological differences between stall flutter and buffeting. Some experimental results are presented concerning both the boundaries at which these phenomena occur and the stresses involved" (p. 1).

Report presenting pressure distributions, tuft patterns, and schileren surveys of a sweptback wing-fuselage combination in the 8-foot transonic tunnel at a range of Mach numbers and angles of attack. The wing had 45 degrees of sweepback, an aspect ratio of 4.0, a taper ratio of 0.6, and an NACA 65A006 airfoil section. Results regarding an angle of attack of 0 degrees, 4 degrees, 6 degrees, 8 degrees, 12 degrees, and 20 degrees are provided.

Report presenting an investigation to determine the low-speed, large-scale characteristics of an airplane configuration with an aspect ratio 2 triangular wing of variable incidence. The model consisted of the variable-incidence wing in combination with a fuselage of fineness ratio 13, a triangular vertical tail, and a thin, unswept, all-movable horizontal tail. Results regarding the lift characteristics of the model with the tail off, longitudinal stability with the horizontal tail on, trim lift and drag characteristics, and directional instability are provided.

"A theoretical method is derived for the determination of the motions and loads during chine-immersed water landings of prismatic bodies. This method makes use of a variation of two-dimensional deflected water mass over the complete range of immersion, modified by a correction for three-dimensional flow. Equations are simplified through omission of the term proportional to the acceleration of the deflected mass for use in calculation of loads on hulls having moderate and heavy beam loading" (p. 1025).

Report presenting the use of wall perforations on supersonic diffusers to avoid the internal contraction-ratio limitation. Experimental results on a preliminary model of a perforated diffuser at Mach number 1.85 are provided. A theoretical discussion of the flow coefficients and the size and spacing of the perforations are included.

The weight breakdowns in bar chart form in the present report have been prepared to give an indication of 'how we spend our weight.' They are expected to focus attention on the design functions that warrant weight reduction activity on future models.