From Introduction: "The purpose of the flight investigation described herein was as a "proof" check of the system and to determine what effect several variables which could not be practicably simulated would have on the operation."

Report presenting the experimental values of damping in roll of the Bell X-1A airplane and its components at 5 Mach numbers at zero angle of attack. The wing was found to be the predominant contributor to the damping in roll throughout the range of Mach numbers, but the contributions of other airplane components were significant in the range of 2.22 to 2.41.

Memorandum presenting an analysis comparing engine performance determined by use of a method involving adjustment of engine cycle calculations using JP-4 fuel with performance obtained using hydrogen as a fuel.

Force and pressure distribution studies to high angles of attack on all-movable triangular and rectangular wings in combination with body at supersonic speeds. The ranges of aspect ratios were, for the triangular wings, 3/8 to 4, and for the rectangular wings, 1 to 3.

Memorandum presenting an experimental investigation conducted at Mach numbers 1.5, 1.8, and 2.0 to determine the effects of several throat boundary-layer bleed configurations on the performance of a 25 degree half-conical side-inlet system. The effects of several flush-slot configurations and a porous-surface bleed were determined over ranges of angle of attack and bleed-duct and main-duct mass flow.

Memorandum presenting an investigation conducted on the size and density of the boric oxide exhaust cloud from a J47-25 turbojet engine operating on trimethyl-borate fuel at sea-level static conditions. Movies and still photographs were taken from the ground and from a helicopter. Results regarding the visual observation and oxide fallout and concentration are provided.

Report presenting an investigation of the effects of changing indentation design Mach number on the aerodynamic characteristics of a 45 degree sweptback-wing-body combination designed for high performance at a range of Mach numbers. Theoretical results of zero-lift wave drag for the wing-body combination agreed well with the experimental data. Results regarding the bodies, systematic series of wing-body combinations, -2 degree angle of incidence, M = 1.4 revised body, and transition are presented.

Report presenting flight testing over a range of Mach numbers to determine the rolling behavior of a swept-wing airplane with lateral-longitudinal coupling. Results regarding the size of the tail area, speed effects, engine gyroscopic effects, and wing-tip extensions are provided.

Report presenting an investigation of a swept-wing fighter airplane in the transonic tunnel to determine store and pylon loads and the effect of store installation of drag and stability, horizontal-tail loads at sideslip angles of 0 degrees and 5 degrees, and the extent of drag-rise reductions possible by enlarging the fuselage afterbody. Results regarding store and pylon installation, horizontal-tail characteristics, and effects of "area-rule" modifications are provided.

Report discussing a flight investigation to obtain experimental information on the handling qualities of a fighter airplane that a human controlled by supplying signals to an attitude type of automatic pilot. The pilots were found to prefer the control-force characteristics provided by a damper rather than a spring feel system. Information about characteristics in pitch, characteristics in roll, dynamic lateral stability, and rough-air flying, tracking, and landing characteristics is provided.

"An investigation was conducted to determine the effect of porous area suction applied to the knees of full-span leading-edge and part-span trailing-edge flaps installed on an unswept-wing airplane model" (p. 1). The results obtained from an angle of attack of 0 degrees and greater than 0 degrees are compared. The effect of deflection, suction requirements, and stability characteristics are described.

An analysis of air refrigeration systems indicated that air cycles are generally less satisfactory than simple heat exchangers unless high component efficiencies and high values of heat-exchanger effectiveness can be obtained. A system employing a mercury-vapor cycle appears to be feasible for refrigerating air that must enter the system at temperature levels of approximately 1500 degrees R, and this cycle is more efficient than the air cycle. Weight of the systems was not considered. The analysis of the systems is presented in a generalized dimensionless form.

Memorandum presenting a low-speed investigation to determine the effect of double-slotted flaps on the aerodynamic characteristics of a 45 degree sweptback wing. The flap had a span of 0.35 wing semispan with the inboard end at 0.16 semispan. Results regarding lift characteristics, pitching-moment characteristics, and drag characteristics are provided.

"Methods are presented for calculating net thrust using air specific-impulse data for various fuels. Nomographic solutions are given to adapt the methods to turbojet-afterburner calculations. These nomographs can be used to compute net thrusts obtained by expanding exhaust gases to either a Mach number of 1.0 or the ambient static pressure at the nozzle exit. Thermodynamic data for several fuels are also presented" (p. 1).

From Summary: "Data are included on the amount of fallout from the cloud and the concentration of boric oxide in the cloud."

The quasi-three-dimensional method for the aerodynamic design of both stator and rotor blades is presented for a turbine having high weight flow and high specific work output. The turbine was designed for a relative critical velocity ratio of 0.82 at the rotor hub inlet. At the condition of equivalent design work output and blade speed, the efficiency based on the overall actual total-pressure ratio was 0.875, which is almost at high as those obtained for conservatively designed turbines.

For use in computing the detailed performance parameters of individual blade rows of a multistage compressor, radial distributions of total pressure, total temperature, static pressure and air-flow angle are tabulated. The data cover a range of air flow from choke to the approximate surge limit for equivalent rotor speeds from 70 to 100 percent of design. Equivalent stage performance curves determined from the radial survey measurements indicate both the individual performance and the matching of the stages at the speeds investigated.

Wind-tunnel tests have been made to determine the static longitudinal stability of several models of a short-range artillery shell at Mach numbers of 0.8, 0.9, 1.0, and 1.2. The results of the tests indicated that the best of the spool-shaped shells was statically stable in pitch at all test Mach numbers for an angle-of-attack range up to about 10 degrees. The best of the finned shells was stable to a maximum angle of attack of about 6 degrees.

Report presenting testing of a model of an airplane configuration with thin, highly tapered, 45 degree sweptback surfaces at a range of Mach numbers to obtain the lateral stability characteristics of the configuration. Results regarding the time history, trim characteristics, lateral-force derivative, vector plots, and downwash and sidewash are provided. At Mach number 1.3, the model experienced a violent combined lateral-longitudinal motion and large angles of sideslip and attack with large rolling velocities.

Report presenting the static-thrust characteristics of the NACA 8.75-(5)(05)-037 dual-rotation propeller with six and eight blades as obtained on the static test stand in the course of an investigation set up primarily for studying blade vibratory stresses. Force data were obtained for a blade-angle range measured at the 0.75 radial station on the front propeller. Results regarding the effect of tip Mach number, effect of number of blades, and effect of blade angle are provided.

Report presenting aerodynamic-heating data obtained on a modified-fineness-ratio-5.0 Von Karman nose shape at free-stream Mach numbers up to 10.4 with a rocket-propelled model. Transient skin temperatures were measured on the nose, and a maximum temperature of 1,663 degrees R was measured after maximum Mach number was obtained.