Report presenting a free-flight investigation of a rocket-powered research model to determine the hinge-moment and normal-force characteristics of a trailing-edge control on a delta wing between Mach numbers of 0.50 and 1.26. The hinge-line location of 40 percent control chord satisfactorily reduced the high hinge moments associated with plain flap-type controls. Results regarding hinge moments and normal force are provided.

Report presenting the overall performance and rotating stall characteristics of a single-stage compressor with a hub-tip ratio of 0.76 and a symmetrical velocity diagram over a wide range of weight flow for several different design speed percentages. No discontinuities in performance were noted at an operating condition. Results regarding the overall performance, types of rotating stall observed, circumferential extent of stall zones, flow fluctuations incited by rotating stall, rotating-stall hysteresis, and a theoretical analysis are provided.

Report presenting an investigation to determine the effect of several seeker-nose configurations on the static longitudinal stability, canard control characteristics, and lift and drag at Mach number 1.60 of a canard-type ramjet missile with 70 degree delta canard control surfaces and 70 degree delta wings. The tests indicate that, for the most part, the configurations tested exhibited no significant difference in static longitudinal stability or horizontal-canard control effectiveness.

From Summary: "Calculated velocity and temperature distributions, as well as relations among Nusselt number, Reynolds number, and friction factor, are presented."

Report investigation Freon-12 as the working fluid for a 14-inch-diameter supersonic mixed-flow axial-discharge rotor of the impulse type. The whole series of tests was performed twice, once with guide vanes that underturned the air and gave a lower relative inlet velocity and once with a set that gave the design values for velocity and direction. Results regarding the overall performance characteristics, entrance flow, internal pressure patterns, exit flow, and rotor losses are provided.

Memorandum presenting an investigation of the use of a small quantity of fluid that ignites spontaneously within nitric acid to initiate the acid-gasoline reaction carried out in a 200-pound-thrust uncooled engine with a characteristic length of 30 inches at a range of temperatures. The starting process was divided into two phases: ignition and transition. Results regarding ignition temperature limits, effect of temperature on allowable igniter fuel flow, optimum valve opening time for ignition, effect of temperature on transition process, effect of running-fuel reactivity on transition, and tar-forming tendencies of igniter fuels are provided.

Report presenting an investigation of an axially symmetric nacelle-type conical spike inlet with two bypasses located in the horizontal plane and on opposite sides of the nacelle in the 8- by 6-foot supersonic tunnel at several Mach numbers and angles of attack. Results regarding the performance with open bypasses and with closed bypasses are provided.

Report presenting an analytical investigation to evaluate the effects of permeability variations on surface temperatures and coolant flows for ranges of temperature and Reynolds number that are of particular interest for gas-turbine blades. Results regarding small-orifice calibrations, effect of orifices in series with porous surfaces on permeability correlation, effect of permeability variations on coolant-flow rates and porous-wall temperatures, evaluation of the experimental porous blade, and design features for transpiration-cooled turbine blades are provided.

From Summary: "An analysis was made to permit the calculation of the effectiveness of oxide coatings in retarding the transient heat flow into turbine blades when the combustion gas temperature of a turbojet engine is suddenly changed. The analysis is checked with experimental data obtained from a turbojet engine whose blades were coated with two different coating materials (silicon dioxide and boric oxide) by adding silicone oil and tributyl borate to the engine fuel. The very thin coatings (approximately 0.001 in.) that formed on the blades produced a negligible effect on the turbine-blade transient temperature response. With the analysis discussed here, it was possible to predict the turbine rotor-blade temperature response with a maximum error of 40 F."

Report presenting an analysis of some longitudinal maneuvers obtained during performance tests of the Convair XF-92A by using measured period and time to damp to half amplitude and with a computer to give a preliminary measurement of the airplane stability and damping at a range of Mach numbers. Results indicated that no loss in control effectiveness was shown, the static stability increased with Mach number, the damping was light but positive, and the damping factor was low.

Report presenting a continuation of the program in the high-speed 7- by 10-foot tunnel to investigate the aerodynamic characteristics in pitch, sideslip, and steady roll of model configurations with variations in wing geometric parameters. This report contains the aerodynamic characteristics in pitch of wing-fuselage combinations with wings of aspect ratio 4, sweepback angle of 45 degrees, and three different taper ratios. Results regarding lift characteristics, pitching-moment characteristics, drag characteristics, and lift-drag ratio are provided.

Report presenting an experimental investigation of a canard-type missile configuration with an underslung scoop inlet in the 8 by 6-foot supersonic wind tunnel at a range of Mach numbers, angles of attack, control-surface deflection angles, boundary-layer-scoop heights, and inlet mass-flow ratios. Measurement of the external forces indicated that maximum lift-drag ratios were in excess of 5 at all test Mach numbers.

Report presenting the results of testing on a high-pressure-ratio, axial-flow compressor rotor designed to operate efficiently at high speeds. High flows are needed to increase engine thrust without increasing engine front area. Results regarding the overall total-pressure ratio and adiabatic efficiency, effect of angle of attack on efficiency, determination of the effect of mean radius rotor inlet Mach number and Reynolds number on turning angle, and other information about ideal operation of the rotor are provided.

Report presenting an investigation of the thrust loading by wake surveys in the slipstream of a full-scale supersonic-type three-blade propeller (the Curtiss-Wright 109622) that was conducted in the transonic tunnel. Integrated-thrust coefficient values were found to be in good agreement with the force data. Results regarding thrust loading at constant advance ratios, at advance ratios for maximum efficiency, and lift-coefficient distribution are also provided.

From Introduction: "This paper presents the results of tests made at a Mach number of 2.01 to determine the effect of body length on the longitudinal characteristics (zero roll angle) for five complete configurations as well as for the bodies alone, the bodies plus wings, and the bodies plus canard surfaces. The experimental results are compared with some simple theoretical estimates."

Report presenting a method for calculating the steady-state rolling effectiveness of flexible sweptback wings at subsonic and supersonic speeds. Comparisons between the experimental and calculated values are presented for a wide range of wing-aileron configurations.

The design principle of injecting liquid fuel at more than one axial station in an annual turbojet combustor was investigated. Fuel was injected into the combustor as much as 5 inches downstream of the primary fuel injectors. Many fuel-injection configurations were examined and the performance results are presented for 11 configurations that best demonstrate the trends in performance obtained. The performance investigations were made at a constant combustor-inlet pressure of 15 inches of mercury absolute and at air flows up to 70 percent higher than values typical of current design practice. At these higher air flows, staging the fuel introduction improved the combustion efficiency considerably over that obtained in the combustor when no fuel staging was employed. At air flows currently encountered in turbojet engines, fuel staging was of minor value. Radial temperature distribution seemed relatively unaffected by the location of fuel-injection stations.

Report presenting free-flight and free-jet testing of a twin side-inlet configuration, including a pilot's canopy and a wheel-well fairing, between a range of Mach numbers from 0.8 to 1.55. Using side inlets provides two important advantages: the inlet can be located closer to the engine and it frees up the nose for other uses. Results regarding the effects of oscillating flow, external-drag coefficient, and total-pressure recovery are provided.

Memorandum presenting an investigation of variable-area auxiliary inlets intended to provide for the variation of engine air requirements with speed and altitude. Three auxiliary scoop inlets were tested in combination with a nose inlet at a range of Mach numbers tested at zero angle of attack.

Report presenting an experimental investigation of pressure disturbances caused by the reflection of two-dimensional shock waves from a slotted boundary at Mach numbers between 1.0 and 1.5. Simple-wedge airfoils with three different apex angles were used to generate shock waves and the static pressures ere measured at various points along the surface of each airfoil.

Report presenting testing of a multiweb wing structure, representing an airplane or missile wing, under simulated supersonic flight conditions to determine the transient temperature distribution. The aerodynamic loads caused the model to fail near the end of the test. The failure is analyzed and it is determined that the model failed as a result of the combined action of aerodynamic heating and loading.

An investigation of secondary-flow loss patterns originating in three sets of turbine nozzle blade passages was conducted by means of flow-visualization studies and detailed flow measurements. For all cases, high loss values were measured in the fluid downstream of the corners formed by the suction surfaces of the blades and the shrouds, and these losses were accompanied by discharge-angle deviations from design values. Despite the size of the loss regions and angle gradients, over-all mass-average blade efficiencies were of the order of 0.99 and 0.98 and, therefore, are not a good index of blade performance.

The effects of inlet circumferential position around the fuselage on the characteristics of a half-conical scoop inlet having a 24.6deg half-angle cone have been investigated in the langley 4- by 4-foot supersonic pressure tunnel. Pressure-recovery results have been obtained at a Mach number of 2.01 for a fixed boundary-layer-bleed height which was 60 percent of the boundary-layer thickness at an angle of attack of 0deg, and for cowling position parameters of 42.4deg and 38.0deg. inlet had a capture area equal to 24.9 percent of the basic-fuselage frontal area. The angle of attack was varied from 0deg to 12deg.

"Aeroelastic instability phenomena of isolated open and closed rigid bodies of revolution free to move under elastic restraint have been investigated experimentally at low speeds by means of models suspended at zero angles of attack and yaw on slender flexible struts from a wind tunnel ceiling. Three types of instability were observed - flutter similar to classical bending-torsion flutter, divergence, and an uncoupled oscillatory instability which consists in nonviolent continuous or intermittent small-amplitude oscillations involving only angular deformations. The speeds at which this oscillatory instability starts were found to be as low as about one-third of the speed at flutter or divergence and to depend on the shape of the body, particularly that of the afterbody, and on the relative location of the elastic axis" (p. 1).