Report presenting an investigation of the performance of a diffuser static-pressure-ratio control with a normal shock-positioning limit on a ramjet engine installed in an altitude free-jet facility. Testing occurred at Mach numbers of 2.35 and 2.50, altitudes of 50,000, 60,000 and 65,000 feet, and angles of attack of 0 and 7 degrees. Results regarding the effects of control constants, effects of disturbance size and operating point, effects of flight conditions, and control limitations and possible improvements are provided.

Injection techniques for fluorine, oxygen, and heptane in 200-pound thrust rocket engine.

"A 1/13-scale model of the forebody of the Republic F-105 with twin-duct wing-root inlets was tested in the Langley 4- by 4-foot supersonic pressure tunnel through a range of angle of attack from -4 deg to 15 deg at a Mach number of 2.01 and a Reynolds number of approximately 3.4 x 10(exp 6) per foot. The tests were made with four configurations which incorporated varying amounts of sweep and stagger of the inlet leading edges, modifications to the areas of the boundary-layer diverter floor plate, and modifications to the area of the boundary-layer diverter bleed slots. The highest overall pressure recovery at an angle of attack of 0 deg (average total-pressure recovery, 0.84 mass-flow ratio, 0.98) was achieved with configuration having an inlet leading-edge sweep angle of 58 deg with no leading-edge stagger" (p. 1).

From Introduction: "The feasibility of closed-loop control of ram-jet-engine thrust is demonstrated in references 1 and 2. The dynamic behavior of a 28-inch-diameter ram-jet engine designed to operate in Mach number range of 2.35 to 2.70 is evaluated herein."

Report presenting an investigation to study some factors affecting the drag of relatively large nonlifting bodies of revolution at transonic speeds in the 8-foot transonic tunnel. Drag and surface pressure measurements were made for two geometrically similar bodies of revolution of 8-inch and 4-inch maximum diameter at zero angle of attack through a range of Mach numbers.

Report presenting a low-speed wind-tunnel investigation to determine the high-lift characteristics of a thin delta wing equipped with a single and double slotted trailing edge flap. Results regarding the lift, drag, and pitching-moment characteristics of the plain wing and various flap configurations are provided. The maximum lift coefficient of the model could be increased from 1.45 for the plain wing to 1.86 when the double slotted flap was deflected to 50 degrees.

Report presenting an experimental investigation to determine some of the characteristics of an inset tab as an aerodynamic balance in the Mach number range between 0.7 and 1.4 using a zero-lift rocket-propelled model in free flight. Results regarding the variation of the aileron floating angle, the wing-tip helix, the balancing effect, and the penalty occurred throughout the Mach number range are provided.

Report presenting an investigation of the hydrodynamic characteristics of a model of a 130,000-pound transonic flying-boat design with a high length-beam ratio planing-tail hull, sweptback gull wings with wing-tip floats, and wing-root jet-power-plant installation. Results regarding the aerodynamic, hydrostatic, and hydrodynamic qualities are provided.

Memorandum presenting an investigation of the stability and controllability in space of an arrangement comprising a man standing on a small platform which is rigidly connected to a jet nozzle with its thrust axis perpendicular to the platform and its thrust opposed to the pull of gravity. Results regarding the initial flights with the platform, flights with landing gear, flights with landing gear and inertia, flights with landing gear and gyroscope, flights with landing gear and control stick, flights with landing gear and seat, and discussion of the balancing process are provided.

Report presenting an investigation in the 7- by 10-foot tunnel to determine the transonic aerodynamic characteristics of a wing with a spanwise variation in thickness ratio. The wings investigated had 45 degrees of sweepback, aspect ratio 8, taper ratio 0.45, and airfoil sections tapered from an NACA 63A010 section at the root chord to an NACA 63A006 section at the tip chord. Results regarding lift characteristics, drag characteristics, and pitching-moment characteristics are provided.

An investigation has been conducted in the Langley 4- by 4-foot supersonic tunnel at Mach numbers 1.40 and 1.59 to determine the effect of aileron profile on the aerodynamic characteristics of a complete model of a supersonic aircraft configuration. The model had a 40 degree sweptback tapered wing with 10-percent-thick circular-arc sections normal to the quarter-chord line. The ailerons were 20-percent chord and were located on the outboard 50 percent of the wing semispans. The various ailerons investigated included the basic circular-arc profile and three flat-sided ailerons having ratios of trailing-edge thickness to hinge-line thickness t of 0, 0.5, and 1.0. Low aileron effectiveness was obtained with the circular-arc and t = 0 profiles. Increasing the trailing-edge thickness (t = 0.5 and 1.0) resulted in increased effectiveness as well as increased hinge moments with only a slight increase in drag.

Memorandum presenting a corrosion investigation conducted with molten sodium flowing at a velocity of 25 feet per second in a toroid of AISI type 347 stainless steel with an insert consisting of seven Globeiron tubes in parallel at the hot zone. A 500-hour test was run at a nominal outer wall temperature of 500 degrees Fahrenheit with a temperature difference between the hot and cold sections of 40 degrees.

Report presenting an investigation of a semispan-wing-fuselage model with a 55 sweptback triangular wing of aspect ratio 3.5 with a horn-balanced flap-type control with a partially inset unbalancing tab. Control hinge-moment and effectiveness characteristics were obtained for a range of angles of attack, flap deflections, and tab deflections.

"Preliminary tests have been made of a small burner to meet the requirements for application to supersonic ram jets. The principal requirements were taken as: (1) efficient combustion in a high-velocity air stream, (2) utilization for combustion of only a small fraction of the air passing through the unit, (3) low resistance to air flow, (4) simple construction, and (5) light weight. Tests of a small burner were carried to stream velocities of nearly 150 feet per second and fuel rates such that one-eighth to one-fourth of the total air was involved in combustion" (p. 1).

Characteristic exhaust velocity over a range of mixture ratios and variations in gas velocity with distance from the injector were measured for six injectors. Comparisons of injector performance showed the gains obtained from oxidant atomization, fuel atomization, and propellant mixing. The results are compared with oxygen and heptane performance and show the effect, which is qualitatively small, of spontaneous propellant ignition on the relation between injection processes and engine performance.

A spike-type nose inlet with sharp-lip cowl was investigated at a free-stream Mach number of 2.5 with water injection in its 16-inch diameter, 11-foot-long subsonic diffuser section. Inlet total temperature of exit with liquid-air ratios of about 0.04 with no apparent change in the critical pressure recovery. The observed temperature drops were less than the theoretically predicted values, and the amount of water evaporated was 35 to 50 percent less than that theoretically possible.

Report presenting power spectral densities of normal load factor obtained for two service operational training flights of a Republic F-84G airplane and three service operational training flights of a North American F-86A airplane in order to indicate the load-factor frequency content and possible uses of power spectral methods in analyzing maneuver load data. Results regarding the spectra of the two different planes, an analytical representation of maneuver load spectrum, and probability distributions are provided.

The performance of a control system designed for variable thrust applications was determined in an altitude free-jet facility at various Mach numbers, altitudes and angles of attack for a wide range of engine operation. The results are presented as transient response characteristics for step disturbances in fuel flow and stability characteristics as a function of control constants and engine operating conditions. The results indicate that the control is capable of successful operation over the range of conditions tested, although variations in engine gains preclude optimum response characteristics at all conditions with fixed control constants.