Report presenting total-pressure recoveries obtained with four-cone inlet combinations at Mach number 1.85. The configurations tested included a cone designed to produce three oblique shocks ahead of the diffuser inlet combined with two other inlets, a cone generated by a parabolic arc in combination with two other inlets, a cone-inlet combination designed to produce an isentropic entrance flow at 0 degrees angle of attack, and a 30 degree single-shock cone combined with a perforated inlet section. Each of the configurations yielded total-pressure recoveries higher than what was reported in previous testing.

Report presenting an investigation of the effect of tip shape on the aerodynamic characteristics of large-scale 45 degrees swept-forward and 45 degrees swept-back wings. The results show that none of the variations of tip shape investigated nor the addition of bodies of revolution at the tips produced major changes in the characteristics of the swept wings.

Report presenting an aerodynamic-control-effectiveness investigation using free-flight rocket-propelled RM-5 test vehicles at high subsonic, transonic, and supersonic speeds. Results regarding aileron control characteristics and drag measurements are provided.

Report presenting calculated performance characteristics of a ducted rocket with gasoline and liquid oxygen as the fuel mixture and with a mass flow of 1 slug per second. Evaluations of net thrust, frontal area, thrust per unit frontal area, and specific fuel consumption are given for all of the flight conditions under which it might operate.

Report discussing an investigation to determine the drag of a configuration with a body of fitness ratio 12 with stabilizing tail surfaces and a 12-percent-thick 30-degree swept-forward wing using the free-fall method. The drag oft he wing and the total drag were measured separately and compared. The swept-forward wing was found to greatly increase the effect of drag on the body-tail combination.

Report presenting an investigation to obtain a survey of the flying characteristics at low speeds of models with low-aspect-ratio delta wings. Four models with triangular plan-form sweptback wings and five models with the tips cut off to give taper ratios were tested. Results regarding the stability and control characteristics of each of the models is provided.

Report presenting flight measurements obtained during acceptance tests conducted by the Bell Aircraft Corporation on the XS-1 research airplane. The primary focus was on the aerodynamic loads, handling qualities, maximum lift, and buffet boundaries of the airplane.

The effects on dynamic lateral stability and controllability of some of the important aerodynamic and mass characteristics are discussed and methods are presented for estimating the various stability parameters to be used in the calculation of the dynamic lateral stability of airplanes with swept and low-aspect-ratio wings.

Report presenting tests in a pressure tunnel to determine the low-speed pitch characteristics of a 42 degree sweptback wing with circular-arc airfoil sections. Information regarding the characteristics of basic wing, leading-edge flap investigation, and effects of fuselage are provided.

Report presenting an experimental study of the pressure fluctuations at the combustion-chamber inlets of a ramjet operating a Mach number of 1.92 in the supersonic tunnel. Results are presented for two burner configurations of different operating characteristics to indicate trends in pressure magnitude, frequency, and wave form. Results regarding the perforated conical flame holder, regenerative-type burner, and diffuser total-pressure recovery are provided.

Report presenting the results of tests to determine the effect of taper on the zero-lift drag of wings of constant exposed aspect ratio at low supersonic speeds. Findings indicated that maximum thickness, leading-edge, and trailing-edge sweep are all important in determining the drag coefficient of a tapered wing.

Memorandum presenting a method for varying the mass flow of supersonic inlets with a circular cross section and a central body. The method consists of changing the size of the entering stream tube by means of an inflatable boot on the surface of the central body.

Report presenting transonic speed testing utilizing a rocket-powered Bell XS-1 airplane with 10-percent-thick wing and 8-percent-thick horizontal tail. Information is provided for flights up to Mach number 1.06 and altitudes of about 40,000 feet. Results regarding the variation of control-surface positions and forces and sideslip angle with Mach number, time history, rudder effectiveness, and lateral oscillations are provided.

Report presenting an investigation of the characteristics of three representative NACA 1-series nose inlets over a Mach number range extending from 0.4 to 0.925 in the high-speed wind tunnel. Results regarding pressure distribution, drag, and some potential design considerations are provided.

Memorandum presenting the results of measurements of gust and draft velocities within thunderstorms at an air field over a 7-day period. The data are summarized in tables and compared to previous flights.

Report presenting various methods of providing compressor-capacity and pressure-ratio control in the gas-generator type of compound engine over a range of altitudes. The analytical results indicated that the best method of control is that in which the first stage of compression is carried out in a variable-speed supercharger driven by a hydraulic slip coupling. Results regarding the constant-pressure-ratio compressor, constant-volume compressor, comparison of piston-type and rotary compressors, effect of variable-area turbine nozzle, and effect of designing for high altitudes are provided.

Report presenting an investigation to determine the effect of the primary-jet temperature on the performance of several ejector configurations. The performance of ejectors expressed in terms of the ratio of weight of secondary air flow to primary air flow was found to be affected by temperature in two ways.

From Summary: "Thrust augmentation of an axial-flow type turbojet engine by burning fuel in the tail pipe has been investigated in the NACA Cleveland altitude wind tunnel. The performance was determined over a range of simulated flight conditions and tail-pipe fuel flows. The engine tail pipe was modified for the investigation to reduce the gas velocity at the inlet of the tail-pipe combustion chamber and to provide an adequate seat for the flame; four such modifications were investigated."

"Spin tests have been conducted in the Langley free-spinning tunnel on a 1/16-scale model of the McDonnell XP-85 airplane with the normal X-tail replaced with a short-coupled conventional tail arrangement. The effect of the conventional tail arrangement and the effects of various modifications upon the spin and recovery characteristics of the model were determined. The results of the tests indicated that installation of the conventional tail arrangement will not provide satisfactory recoveries from spins of the airplane" (p. 1).

"Low Mach number longitudinal-stability and control characteristics as predicted by use of wind tunnel data from a powered 3/16-scale model are compared with flight test measurements of a Navy BTD-1 airplane. The accuracy of the wind tunnel data and the discrepancies involved in attempting to correlate with flight data are discussed and analyzed. The comparison showed that wind tunnel predictions were, in general, in good agreement with flight test data" (p. 1).

From Summary: "The hottest part of a turbine blade is likely to be the trailing portion. When the blades are cooled and when water is used as the coolant, the cooling passages are placed as close as possible to the trailing edge in order to cool this portion. In some cases, however, the trailing portion of the blade is so narrow, for aerodynamic reasons, that water passages cannot be located very near the trailing edge. Because ceramic coatings offer the possibility of protection for the trailing part of such narrow blades, a theoretical study has been made of the cooling effect of a ceramic coating on: (1) the blade-metal temperature when the gas temperature is unchanged, and (2) the gas temperature when the metal temperature is unchanged."

"As part of an investigation of the performance and operational characteristics of the TG-100A gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100R. The highest compressor pressure ratio was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475R" (p. 1).

An altitude-wind-tunnel investigation of a TG-100A gas turbine-propeller engine was performed. Pressure and temperature data were obtained at altitudes from 5000 to 35000 feet, compressor inlet ram-pressure ratios from 1.00 to 1.17, and engine speeds from 800 to 13000 rpm. The effect of engine speed, shaft horsepower, and compressor-inlet ram-pressure ratio on pressure and temperature distribution at each measuring station are presented graphically.

"Tests of two 10-foot-diameter two-blade propellers which differed only in shank design have been made in the Langley 16-foot high-speed tunnel. The propellers are designated by their blade design numbers, NACA 10-(5)(08)-03, which had aerodynamically efficient airfoil shank sections, and NACA 10-(5)(08)-03R which had thick cylindrical shank sections typical of conventional blades. The propellers mere tested on a 2000-horsepower dynamometer through a range of blade-angles from 20 degrees to 55 degrees at various rotational speeds and at airspeeds up to 496 miles per hour" (p. 1).