"Eighty pounds of butylsilanes required for test purposes was synthesized by reducing butyltrichlorosilane with lithium aluminum hydride in dioxane solution. This was safely and successfully accomplished by operating in steel equipment under an atmosphere of oil-pumped nitrogen to minimize the hazards involved in conducting this operation on a large scale. The apparatus and procedure used to carry out the necessary manipulations are described in this report" (p. 1).

Report presenting testing of gaseous hydrogen fuel burned in a connected-pipe combustor with a cross section equal to 35 degrees sector of a 28-inch diameter. Eleven shrouded fuel-injector configurations were used to obtain combustion data at specified high-altitude ramjet combustor conditions. Results regarding shroud air blockage, mixing tabs, higher equivalence-ratio burning, injector size, flameholder durability, burner length, combustion efficiency, and temperature profile are provided.

Report presenting an investigation of the hydrodynamic take-off and landing characteristics of a model of a hypothetical jet- and rocket-propelled high-speed airplane fitted with various designs of a single monoplane hydrofoil mounted near the center of gravity. This testing is performed to determine the feasibility of the water-based operation of high-speed airplanes. Results regarding the general hydrodynamic characteristics, development of hydrofoils, and effects of some hydrofoil parameters are provided.

Report presenting the results of an experimental investigation of the dynamic stability and control of a simple ducted-fan model in hovering flight in order to provide basic information on the stability and control of jet vertically rising airplanes in hovering flight. The investigation primarily consisted of flight tests with the model hovering at altitude and near the ground.

The factors affecting the performance of ejector suitable for aircraft cooling are investigated theoretically and experimentally. The investigation covers a range of shroud-to-nozzle diameter ratios from 1.1 to 1.6, of shroud lengths from 0.2 to 2.28 nozzle diameters, of secondary-to-primary weight-flow ratios from 0 to 0.12, and of ambient-to-nozzle pressure ratios from 0.7 to 0.06. The results of a simplified theoretical analysis based on each type of flow are in good agreement with those experimentally obtained.

Report presenting an investigation of turbine operating temperatures as affected by liquid cooling, especially if materials of high conductivity are used, at gas temperatures up to 1925 degrees Fahrenheit. Results regarding the statio heat-transfer rig and turbine rig are provided. The investigation showed that nonstrategic material such as aluminum alloy can be used in liquid-cooled turbine wheels at high gas temperatures.

Report presenting wind-tunnel measurements of the pressure distribution on a semispan wing with a modified NACA 0012-64 airfoil section and a plain trailing-edge aileron with the wing unswept and also swept back 45 degrees. Adverse changes in loading and loss of aileron effectiveness developed with the wing unswept as the Mach number was increased, but did not occur at Mach numbers up to 0.925 with the swept wing. Results regarding section characteristics and effect of fences are provided.

Memorandum presenting an investigation of modifications to an indented body which was originally designed on the basis of the transonic drag-rise rule to determine the effects on the aerodynamic characteristics of a 45 degree sweptback-wing-body combination. The investigation covered a range of Mach numbers and angles of attack. Results regarding lift, drag for the nonlifting case, drag for the lifting case, maximum lift-to-drag ratios, and pitching moment are provided.

From Summary: "A procedure is presented for determining boundary-layer quantities from pressure-rake data, which include the combined effects of viscous and shock losses. The problem is analyzed using schileren photographs of the shock configuration, the continuity of mass relationship, and the characteristic of the turbulent boundary layer that its outer edge is defined by a rapid change in slope in the Mach number profiles in the vicinity of the edge."

"In the present paper attention is directed to the aerodynamic parameters, the so-called stability derivatives, that affect the lateral behavior of airplanes and missiles. The discussion is centered on three important quantities, the effective-dihedral derivative, the directional-stability derivative, and the damping-in-roll derivative. These quantities are considered for a large angle-of-attack range at subsonic speeds" (p. 1).

From Summary: "Drag measurements at transonic speeds on rectangular airfoils and on airfoils swept back 450 are reported. These airfoils, which were mounted on cylindrical test bodies, are part of a series being tested in free drops from high altitude to determine the effect of variation of basic airfoil parameters on airfoil drag characteristics at transonic speeds. These rectangular and swept-back airfoils had the same span, airfoil section (NACA 65-009), and chord perpendicular to the leading edge. The tests were made to compare the drag of rectangular and sweptback airfoils at a higher aspect ratio than had been used in a similar comparison reported previously."

Tests have been conducted to determine the drive-motor torque and the static force and moment characteristics of the AN/SPS-1 radar antenna. Shifting the longitudinal position of the antenna had very little effect on the drive-motor torque, which reached a maximum value expressed in terms of dynamic pressure (T/q)(sub max) of 1.15 at an azimuth angle of 245. The maximum observed values of rolling, pitching, and yawing moments in terms of dynamic pressure are -29.0, 66.6, and 13.4, respectively.

Report presenting results from a limited flight program to obtain information on airplane response and actual rates and amounts of control motion used by service pilot in performance of squadron operational training. Results regarding control deflection, control rates, airplane angular velocities, airplane angular accelerations, angle of attack and angle of sideslip, airplane load factors, and airspeed and altitude are provided.

Report presenting the results of deflection tests on the wing of a full-scale, swept-wing jet bomber in the form of structural influence coefficients relating the deflection of a system of points on the wing and concentrated loads applied on the wing spars. The procedures for determining the coefficients are included.

Report presenting some measurements of the vortex movements with time about an airfoil undergoing a blast of sufficient strength to exceed the stall angle by a large amount. Measurements were also obtained without flight simulation but with blast orientation and strength and compared to the first test. Results regarding Schileren photographs, presentation of results, and discussion of results are provided.

Testing occurred using the NACA wing-flow method at a range of Mach numbers to determine the effect on aerodynamic characteristics of reversing a triangular wing with a 6-percent-thick biconvex section. Results regarding the lift characteristics, drag characteristics, and pitching-moment characteristics are provided.

Memorandum presenting a method for calculating the stick-fixed longitudinal stability of a wing-fuselage configuration at subcritical Mach numbers. The method applies to unswept- and swept-wing configurations. The stability parameters estimated by the method show reasonable agreement with the experimental values for the 23 configurations used in the comparison.

Report presenting the effect of adjusting compressor inlet-guide-vane angle on overall compressor performance, stage performance, engine thrust, and specific fuel consumption in a 7000-pound-thrust turbojet engine. Data was obtained at several angles and demonstrated that increasing the guide-vane angle decreased the weight flow at the higher compressor speeds. It had no effect on stage performance other than at the first stage, where it reduced maximum flow, pressure, and temperature coefficients.

The stator-blade angles in the twelfth to fifteenth stages of a 16-stage high-pressure-ratio axial-flow compressor were decreased 3 deg The over-all performance of this compressor is compared with the performance of the same compressor with standard blade angles. The matching characteristics of the modified compressor and a two-stage turbine were also obtained and compared with those of the compressor with the original blade angles and the same turbine.

Report presenting an investigation of the certain wing-body combinations as measured by the lift-drag ratio for two arrangements: a half-cone mounted under a triangular wing and a half Newtonian ogive similarly mounted. The configurations were tested for a range of Mach numbers from 1.5 to 5.0. Results regarding the comparison of a half-cone versus a full cone and calculated maximum lift-drag ratio are provided.

Report presenting a two-stage air-cooled turbine configuration and its velocity diagrams for use in the design of blade shapes for a set of engine conditions suitable for flight at Mach number 2.5. The method used to determine the velocity diagrams was one in which the turbine frontal areas and rotor hub inlet and outlet velocities were minimized by varying the hub-tip ratio, turbine- to compressor-tip-diameter ratio, and the work split between the first and second stages.

A blade design study is presented for a two-stage air-cooled turbine suitable for flight at a Mach number of 2.5 for which velocity diagrams have been previously obtained. The detailed procedure used in the design of the blades is given. In addition, the design blade shapes, surface velocity distributions, inner and outer wall contours, and other design data are presented. Of all the blade rows, the first-stage rotor has the highest solidity, with a value of 2.289 at the mean section. The second-stage stator also had a high mean-section solidity of 1.927, mainly because of its high inlet whirl. The second-stage rotor has the highest value of the suction-surface diffusion parameter, with a value of 0.151. All other blade rows have values for this parameter under 0.100.

Report presenting an experimental investigation to determine the vibration phenomena that occur in turbine blades of a typical jet-propulsion engine during service operation; high-temperature strain gages were used to measure the turbine-blade vibrations. Results regarding modes of vibration, centrifugal stresses, and vibration of turbine blades during operation are provided.

Report presenting a J47-25 engine instrumented with thermocouples and operated under steady-state and transient-conditions in an altitude test stand to determine the operating temperatures at altitude in the turbine section of a turbojet engine. Results regarding the temperature effects and transient operation are provided.