From foreword: The TRIGA is a novel and unusually safe training, research, and isotope-production reactor that contains solid homogeneous uranium-zirconium-hydride-moderator fuel elements and operates at power levels in the range of from 10 kw to 250 kw.

From Introduction: "The overall performance and operational characteristics are reported in reference 1. The performance of the components (compressor, combustor, turbine) is reported herein."

Report presenting an investigation in the supersonic wind tunnel to evaluate the performance of a 16-inch ramjet engine using a can-type-combustor configuration. Data were obtained over a range of fuel-air ratios, free-stream Mach numbers, and angles of attack. Combustion efficiencies were generally unaffected by free-stream Mach number and varied between 60 and 90 percent with JP-4 fuel.

Memorandum presenting an investigation of the influence of combustion-chamber length on afterburner performance under conditions simulating altitude flight. Results regarding afterburner-inlet conditions, operational limits, and structural cooling requirements are provided.

"A rocket-propelled model of the X-3 configuration equipped with an all-movable tail of aspect ratio 3.0 has been flown to determine the longitudinal characteristics of this configuration at high lift coefficients. An analysis of the response of the model to rapid deflections of the horizontal tail gave information on lift, drag, longitudinal stability, and longitudinal trim change. The primary result of the tests was that the configuration was indicated to have very unstable tendencies at lift coefficients above the stall and at Mach number near 0.7" (p. 1).

Report presenting a wind-tunnel investigation to determine the low-speed static and rotary stability derivatives of a model of the Douglas D-558-II airplane in the landing configuration. Results regarding the static longitudinal characteristics, static lateral characteristics, characteristics in rolling flow, and characteristics in yawing flow are provided.

"An investigation of two semispan wings swept back 50 degrees 38 minutes was conducted in the Langley high-speed 7- by 10-foot tunnel over a Mach number range of 0.68 to 1.15 by use of the transonic-bump technique. This paper presents the results of the investigation of wing-alone and wing-fuselage configurations of the two wings; one was an untwisted uncambered wing and the other was the same wing but incorporated twist and camber designed to give uniform load at a lift coefficient of 0.25 at a Mach number of 1.10. The semispan wings had their quarter-chord lines swept back 50 degrees 38 minutes, aspect ratios 2.98, taper ratios 0.45, and modified NACA 64A-series airfoil sections tapered in thickness" (p. 1).

Report presenting the use of the matrix-integration method for calculating static aeroelastic phenomena on a wing with concentrated aerodynamic forces at the tip due to tip tanks or boom-mounted lifting surfaces. The results indicate that the presence of a tip tank on an unswept wing tends to deteriorate its static aeroelastic characteristics and that a lift surface geared to the aileron and mounted on a boom ahead of the tip may improve the static aeroelastic characteristics to warrant consideration of a vane as a device for relieving adverse aeroelastic effects.

Report presenting the results of strength tests on thick-skin wing structures. Some of the factors that were found to have influence on the buckling behavior and strength are shown with data and the conditions under which a combination of multiweb and multipost construction may be used are discussed. Results regarding circular-arc airfoils, multiweb beams, and multipost stiffened beams are provided.

Report presenting a flight investigation of the aerodynamic loads of a partial-span extended leading-edge slat on the 35-degree sweptback wing of the Douglas D-558-II research airplane. The slat normal-force coefficient was found to increase with airplane normal-force coefficient until peak slat normal-force-coefficient was reached at approximately 2.2 to 2.4.

Report presenting the development and use of a chart for use in estimating the combustion efficiency of a turbojet at altitude flight conditions. The combustion efficiency chart is based on the assumption that the corrected gas flows, pressures, and temperatures of engine components other than the combustor are unique functions of corrected engine speed.

From Introduction: "The combustor efficiency, combustor-inlet Mach number, and combustor total-pressure ratio of the annular-piloted combustor are presented herein. The effect of combustor-exit total pressure on combustor efficiency, and a comparison of configurations on the basis of specific fuel consumption are also presented."

Report presenting an investigation of several double-ramp inlets utilizing a variable-angle second ramp mounted on the fuselage of a supersonic airplane with a twin-duct air intake system and investigated at several Mach numbers. With all of the inlets, the boundary-layer air from the precompression ramp bridged across the leading edge of the variable ramp.

Discussing the petrography of the black Toldilto Limestone Area with respect to the processing properties

An improved continuous ether extractor is described. The modifications include a means of safely disposing of the active raffinate and a means of positivity checking the raffinate for completeness of uranium extraction. The results obtained on synthetic samples and on dissolver solutions are given. This work was undertaken because of a need for the determination of uranium in dissolver solutions with an accuracy of 0.1 per cent. After a review of available methods it was decided that a gravimetric determination would meet the requirements of precision and accuracy.

Report presenting an investigation of the influence of combustion-chamber length on afterburner performance under conditions simulating altitude flight. The component design and arrangement were selected from the results of previous investigations to produce maximum thrust augmentation near the stoichiomtric condition. Results regarding afterburner-inlet conditions, operational limits, and structural cooling requirements are provided.

Four solid-propellant rocket engines of nominal 1000-pound-thrust were tested for starting characteristics at pressure altitudes ranging from 112,500 to 123,000 feet and at a temperature of -75 F. All engines ignited and operated successfully. Average chamber pressures ranged from 1060 to ll90 pounds per square inch absolute with action times from 1.51 to 1.64 seconds and ignition delays from 0.070 t o approximately 0.088 second. The chamber pressures and action times were near the specifications, but the ignition delay was almost twice the specified value of 0.040 second.

The results of this paper show that a variation is sensing voltage produces an error in timing which increases logarithmically with time interval. Also, variation in supply voltage produces the same type of errors, but the errors are decreased by the ratio of the sensing to supply voltage. An interesting conclusion from this analysis is that if the ratio of supply to sensing voltages is maintained constant, no change in timing will result.