Report presenting an investigation to determine the effects of number of wings on the aileron rolling effectiveness of an RM-5 test vehicle using the free-rolling wind-tunnel testing technique through a speed range to Mach number 0.9. The wings tested had neither taper nor sweepback and were equipped with full-span 20-percent-chord sealed and faired ailerons. Results indicated that increasing the number of wings resulted in a decrease in rolling effectiveness so that results obtained from the test of the three-wing model were closer to conventional-airplane-configuration results than the four-wing data.

Report presenting the performance of a ramjet unit consisting of an intake diffuser, an exhaust nozzle, and a cluster of thin-plate burners contained in a semicircular combustion chamber as investigated in the induction aerodynamics laboratory.

"An investigation of the stability and control characteristics of a 1/10-scale model of a Canadian tailless glider has been conducted in the 10 Langley free-flight tunnel. The glider designated the N.R.L. tailless glider has a straight center section and outboard panels sweptback 43 deg. along the leading edge of the wing. The aspect ratio is 5.83 and the taper ratio is 0.323. From the results of the investigation and on the basis of comparison with higher-scale static tests of the National Research Council of Canada, it is expected that the longitudinal stability of the airplane will be satisfactory with flap up but unsatisfactory near the stall with flap down" (p. 1).

Report presenting a low-speed investigation in the stability tunnel to determine the effects of changes in profile and aspect ratio on the low-speed static-stability and rolling characteristics of triangular wings. The investigation was expanded to determine the effects of adding fins to the upper surface and of cutting portions from the tips of a triangular wing to form low-aspect-ratio tapered wings.

Report presenting the planing characteristics, as determined by tank tests, of three surfaces representative of hydro-ski forms. One surface had a rectangular plan form with a flat bottom, one had a rectangular plan form with a transversely curved bottom, and the third had a flat bottom and triangular plan form. Results are provided in the form of plots of the load on the water, resistance, trimming moment, and draft against total wetted area with speed and trim as parameters.

From Introduction: "An analytical method for estimating turbine performance from the blade angle8 and flow areas was therefore developed at the NACA Lewis laboratory and is described herein."

An investigation is being conducted to determine the feasibility of studying aileron buzz by means of the wing-flow method. Two semispan models which had an aspect ratio of 6 and a taper ratio of 2 with quarter-chord half-span mass-balanced ailerons have been used. One had an NACA 23012 airfoil section and the second, an NACA 65-212 airfoil section. The ailerons on both models were subject to buzz over a small range of Mach number near 0.9. Data obtained by wing-flow testing agreed reasonably well with full-scale flight results.

"Data obtained in tests of a 10-foot diameter, three-blade propeller, having NACA 10-(3)(08)-03 blades, conducted in the Langley 16-foot high-speed tunnel are presented. The propeller performance quantities related by the tests are thrust, torque, efficiency, and advance ratio for various rotational speeds or stream Mach numbers with blade angle as a parameter. Advance Mach numbers varied from 0.12 to 0.64" (p. 1).

Report presenting the results of pressure-distribution measurements obtained in flight over the free-blown canopy of a fighter-type airplane. The measurements were compared to those of the same canopy in wind-tunnel testing to determine the degree of correlation and effects of Mach number and distortion on the pressure distribution.

Report presenting a wing-tunnel investigation in straight and pitching flow to determine the effects of independently varying aspect ratio and angle of sweep on the longitudinal rotary stability characteristics of a series of ten untapered wings. The investigation showed that an interdependent relationship existed between the effects of aspect ratio and sweep. Results regarding pitching moment due to pitching velocity and lift due to pitching velocity are provided.

Memorandum presenting an investigation in the free-flight tunnel to determine the effect of the sloshing of fuel in partly filled, unbaffled tanks on the lateral stability of a free-flying model. Flight tests were made to determine the effects of the water sloshing for different depths of water and various masses and moments of inertia of the model. The sloshing caused small-amplitude, high-frequency lateral oscillations which were superimposed on the normal Dutch roll oscillation so that the lateral motions of the model appeared jerky.

Report presenting testing on the first of a series of flutter rockets, designated the NACA FR-1-A, which was tested with two identical swept wings. Results regarding the launching, flight, and wing failure are provided.

An investigation conducted in the icing research tunnel on a ramp type recessed fuel cell vent assembly to determine its aerodynamic, rain, and icing characteristics. Although the vent-tube openings are located in the region of maximum ramp pressure, vent-tube static pressures are marginal for a low flight speed.

"This paper presents the results of flight tests to determine the lateral and directional stability and control characteristics of the Grumman F8F-1 airplane with three vertical-tail configurations. The data presented herein have no bearing on the performance characteristics of the airplane, which were not measured but which were considered to be exceptionally good. The conclusions reached regarding the lateral and directional stability and control characteristics are summarized in this report" (p. 1).

Memorandum presenting a comparison between the time history of skin temperature measured on the nose of a V-2 and the temperature computed using Eber's experimental relation for heat-transfer coefficients for conical bodies under supersonic conditions. A general method developed for making skin-temperature calculations is used to compute the variation of skin temperature with time for a wide range of values of the pertinent parameters. Results regarding the validity of skin-temperature calculations, ranges of skin temperature, and skin temperatures for a typical supersonic airplane are provided.

"The flight investigation of the C-54D airplane was initiated to determine the necessity of changes or additions to existing handling-qualities requirements to cove the case of instrument approaches with large airplanes. This paper gives a brief synopsis of the results and presents the measured data of tests to determine the stability and control characteristics. It was found that no new requirements were necessary to cover the problems of instrument approaches" (p. 1).

Memorandum presenting an investigation of thrust augmentation by the injection of water-alcohol mixtures into the compressor inlets of a turbojet engine with a centrifugal-flow-type compressor at altitudes of sea level, 5000 feet, and 10,000 feet. The investigation was made to determine the water-alcohol mixture and the injection rate for optimum thrust augmentation. At a standard NACA altitude of 10,000 feet and an engine speed of 16,000 rpm, the mixture and injection rate for optimum thrust augmentation was found to be 20-percent alcohol to water by weight injected at a rate of approximately 1.45 pounds per second.

From Summary: "An investigation was made by the NACA wing-flow method to determine the longitudinal stability and control characteristics at transonic speeds of a semispan model of the XF7U-1 tailless airplane. The 25-percent chord line of the wing of the model was swept back 35 deg. The airfoil sections of the wing perpendicular to the 25-percent chord line were 12 percent thick. Measurements were made of the normal force and pitching moment through an angle-of-attack range from about -3 deg to 14 deg for several ailavator deflections at Mach numbers from 0.65 to about 1.08."

"The results of an experimental investigation made for the purpose of developing suitable jet-engine nacelle designs for a high-speed medium bomber are presented. Two types of nacelles were investigated, the first enclosing two 4000-pounds-thrust jet engines and a 65-inch-diameter landing wheel and the second enclosing a single 4000-pounds-thrust jet engine. Both types of nacelles were tested in positions underslung beneath the wing and centrally located on the wing" (p. 1).

Tests have been conducted in the Langley high speed 7- by 10-foot tunnel over a Mach number range from 0.40 to 0.91 to determine the stability and control characteristics of an 0,08-scale model of the Chance Vought XF7U-1 airplane. The longitudinal-control characteristics of the complete model are presented in the present report with a limited analysis of the results.

Report discusses the development and testing of a rocket-propelled test vehicle to investigate aerodynamic control effectiveness at high subsonic, transonic, and supersonic speeds. Modifications to the section thickness and wing sweepback to improve performance are also described. A description of the vehicle, instrumentation, accuracy, and evaluation of testing results is included.

Memorandum presenting an investigation of the effect of fuel composition and boiling point on the performance of the type of I-16 jet-propulsion engine. Testing occurred with 14 fuels embodying different types of hydrocarbon and with a range of boiling points. The results indicated that fuel composition and boiling range have a negligible effect on engine thrust, rotor speed, and gas temperatures for the principal types of hydrocarbon fuel when used for short periods of time.

Report presenting a study of the effect of fuel properties on carbon deposition in jet-propulsion engine combustors using seven fuels: kerosene, Diesel fuel oil, toluene, xylene, 62-octane gasoline, a commercial solvent, and AN-F-32 (JP-1) in an I-16 engine at static sea-level conditions and constant rotor speed. Results regarding the reproducibility of data, reduced exhaust-jet-nozzle area, and fuel comparison are provided.

An analysis has been made of the characteristics of several cooling cycles suitable for cockpit cooling of supersonic aircraft. All the cycles considered utilize the difference between dynamic and ambient static pressure to actuate the cooling system and require no additional power source. The results of the study indicate that as flight speeds become greater, increasingly complex systems are required to reduce the altitudes above approximately 35,000 feet, a system composed of an externally loaded expansion turbine in conjunction with a supersonic diffuser would maintain tolerable ventilating air temperature, at least up to a flight Mach number of 2. The most complex system considered,composed of compressor, intercooler, and expansion turbine with the intercooler cooling air decreased in temperature by expansion through an auxiliary turbine is capable of maintaining a ventilation air temperature less than ambient temperature up to a flight Mach number of 3.7.