Report presenting an experimental investigation of a high subsonic Mach number turbine with high suction-surface diffusion. The subject turbine was designed for a high weight flow per unit frontal area, a high specific work output, and a relative critical velocity ratio of 0.82 at the rotor hub inlet.

The acceleration characteristics, in the region of maximum acceleration and compressor stall and surge, of an axial-flow turbojet engine with a fixed-area exhaust nozzle were determined by subjecting the engine to fuel flow steps, ramps, and ramps with a sine wave superimposed. From the data obtained, the effectiveness of an optimalizer type of control for this engine was evaluated. At all speeds above 40 percent of rated, a maximum acceleration was not obtained until the engine reached the point of stall or surge. A sharp drop, as high as 80 percent of maximum, in acceleration then occurred as the compressor entered surge of stall. With the maximum acceleration occurring at the point of surge or stall, the optimalizer-type control could not prevent the engine from entering surge or stall. Effective operation of the control may still be possible by sensing the sharp drop in acceleration experienced at the point of stall or surge and using this signal to limit fuel flow. The success of this type of operation would depend on the magnitude of the stall-recovery hysteresis.

Report presenting temperature data for air-cooled turbine rotor blades obtained during an experimental investigation conducted in an altitude test chamber to determine some of the problems pertinent to the operation of air-cooled turbojet engines at a range of turbine-inlet temperatures. Results regarding local blade temperatures, effect of blade-inlet cooling-air temperature on average rotor blade temperature with two types of cooling, effect of cooling-air flow on average rotor blade temperature for a range of turbine-inlet temperatures, cooling-air flow requirements, and blade temperature correlation are provided.

Memorandum presenting an axial-flow turbojet engine operated on a sea-level static test stand to determine whether or not detectable signals were present in pressures and blade stresses that are usable as stall warnings in control applications. Surge and stall behavior were closely examined for the unique characteristics that could be used to control engine behavior in the vicinity of surge and stall. Results regarding a steady-state survey, acceleration study, investigation of an optimalizing technique, and an examination of engine damage are provided.

Memorandum presenting an investigation of the problem of performance losses occurring when airplanes equipped with afterburner and cooling-air ejector installations are flown with the afterburner inoperative. Two sources of the performance losses occurring under some operation conditions were identified: an overexpansion of the propulsive jet and an internal shock system and the excessive pumping action resulting from the off-design.

The forced-convection heat-transfer characteristics of lead-bismuth eutectic were experimentally investigated. Experimental values of Nusselt number for lead-bismuth fell considerably below predicted values. The addition of a wetting agent did not change the heat transfer characteristics.

Memorandum presenting an investigation conducted in the icing research tunnel to determine the icing characteristics, the effects of icing on radar operation, and the protection requirements for two radome configurations. The radomes were for the F-89 airplane and were investigated at airspeeds up to 290 miles per hour, air total temperatures from -15 to 20 degrees Fahrenheit, water contents up to 1.0 gram per cubic meter, and angles of attack of 0 and 4 degrees.

"An experimental investigation of the stage performance of the J71 three-stage turbine at design speed and pressure ratio was conducted as part of a program involving research on the design and operational characteristics of high-work-output multistage axial-flow turbines. The investigation indicated that the over-all work output of the turbine based on temperature measurements at design speed and pressure ratio was 95 percent of the design value, with the greatest loss in work output and efficiency occurring in the first stage. Choking in the second-stage stator over most of the pressure ratios investigated prevented the attainment of design first-stage work output and also prevented any increase in work output of the first stage with increasing over-all pressure ratio at design speed" (p. 1).

Report presenting testing of three wings of different aspect ratios at a range of angles of attack to provide information on the effects of large angles of attack on the lift and normal force on triangular wings in a specific Mach number range. Results regarding the comparisons with data from semispan tests, prediction of lift and normal force and comparison with the experiment, and maximum lift coefficient are provided.

Engine tests were made on a transpiration-cooled blade that was fabricated from an internal load-carrying member with an external surface of wire cloth. After operation in the engine, some damage was noted at the tip region of the trailing edge of the blades. On other sections of the blade, the wire cloth did not appear greatly overheated, and it appeared that satisfactory chordwise temperature distribution was provided by orifices in the blade base.

Memorandum presenting a determination of the external-shock drags of supersonic inlets with circular cross sections from shadow photographs at a Mach number of 2.70 for a low-external-compression and a high-external-compression inlet. The calculated mass flow for both types of inlets showed reasonably good agreement with the measured values. Results as determined by the characteristics method and the approximate method are provided.

Report discussing the cooling characteristics of a 14-cylinder double-row radial air-cooled engine in a four-engine airplane. The effects of charge-air flow, cooling-air pressure drop, and fuel-air ratio on the cooling characteristics were measured separately. The cooling equation, rear middle-barrel temperature, cooling-limited manifold pressure, and maximum cruising power versus temperature-limited power are described.

Report presenting the X-3 airplane, which has a straight 4.5-percent-thick wing of modified hexagonal section, which has been flown to maximum wing normal-force coefficients in the Mach number range from 0.7 to 1.1 at an average altitude of 30,000 feet. Measurements were made of airplane and wing-panel maximum normal-force coefficients and of some buffeting characteristics.

"During the flight program on the Bell X-5 airplane with 59 degrees sweepback to determine the practical Mach number and normal-force coefficient limits of this configuration, a reduction in static longitudinal stability was encountered in maneuvering flight. A determination of the boundary for reduction of longitudinal stability extending to a Mach number of 0.98 is presented in this paper. A reduction of static longitudinal stability existed for all elevator and all stabilizer-executed maneuvers" (p. 1).

Report presenting lift and drag characteristics of the Convair YF-102 airplane for symmetrical wing configuration and for the cambered wing configuration. Data were obtained for a variety of lift coefficients, altitudes, and Mach numbers. Results regarding lift, drag, and some flight-tunnel comparisons are provided.

Report presenting a flight investigation using rocket-powered models to determine some of the effects of wing-to-tail distance on the stability and control characteristics of a canard-missile configuration with 60 degree delta wings and control surfaces. Both models were tested over a range of Mach numbers from 0.8 to 2.0, and canard hinge moment and model drag data are presented. Testing indicated that additional fuselage length increases the lift-curve slope slightly and has very little effect on minimum drag.

Memorandum presenting flight tests with a fighter airplane to study the effect of boundary-layer suction aft of the shock wave on airfoil drag at supercritical Mach numbers and high Reynolds numbers. Airfoil chord force was determined from pressure-distribution measurements obtained at a range of Mach numbers at steady dives. Results of the tests showed no measurable effect of suction for the suction coefficient available.

Report presenting flight measurements of the lateral control characteristics of the Douglas D-558-II airplane in abrupt rudder-fixed aileron rolls. Results regarding the aileron rolling effectiveness, time histories, and aileron effectiveness variation with Mach number are provided.

Report presenting testing of a low-acceleration transonic flutter test vehicle to obtain flutter data on two similar sweptback wings which indicated that wing flutter was symmetrical in mode. Results regarding flight and flutter characteristics for the FR-1-B are provided.

"A flight test of the Aero jet Engineering Corporation's 7KS-6000 T-27 Jato rocket motor was conducted at the Langley Pilotless Aircraft Research Station at Wallops Island, Va, to determine the flight performance characteristics of the motor. The flight test imposed an absolute longitudinal acceleration of 9.8 g upon the rocket motor at 2.8 seconds after launching. The total impulse developed by the motor was 43,400 pound-seconds, and the thrusting time was 7.58 seconds" (p. 1).

"The magnitudes of the yawing moments produced by various parts of an airplane during spins have previously been found to be of major importance in determining the nature of the spin. Discrepancies in resultant yawing moments determined from model and full-scale tests, however, have indicated the probable importance of scale effect on the model. In order to obtain data for a more detailed comparison between full-scale and model results, flight tests were made to determine the yawing moments contributed by various parts of an airplane in spins. The inertia moment was determined by the usual measurement of the spinning motion, and the aerodynamic yawing moments on the fuselage, fin, and rudder were determined by pressure-distribution measurements over these parts of the airplane" (p. 305).

Memorandum presenting two rocket-propelled models flight tested at Mach numbers from 0.8 to 1.5 to determine the effect of a fuselage indentation on the drag of a 52.5 degree sweptback-wing-body configuration with two large strut-mounted store symmetrically located above and below the fuselage. The fuselage had a symmetrical, Mach number 1.0 indentation designed to cancel the wing area normal to the plane of symmetry. Results regarding total drag and pressure drag are provided.

Memorandum presenting a flight investigation to determine the effect of a propulsive jet on the zero-lift drag characteristics of two twin-exit boat-tailed bodies at transonic speeds. The two models had ratios of jet area to base area of 0.394 and 0.590 and covered a Mach number range of 0.8 to 1.15. A slight reduction in drag coefficients from power-off values are obtained during power-on flight for both models.

Report presenting a simplified method for obtaining free-flight measurements of damping in roll through the use of rocket-powered models. Initial configurations have been tested for a range of Mach numbers. Results regarding the rolling velocity with two different airfoil sections and damping-in-roll coefficient are provided.