An evaluation of the effect of geometric design variables on the performance of a 16-inch-diameter ram-jet combustor total pressures from 7 to 44 inches of mercury absolute. Equivalent ratio was varied from 0.1 to 1.0 at a combustor inlet temperature of 1100 degrees R.

Report presenting an experimental determination of the flow fields near swept- and unswept-wing-fuselage combinations at zero sideslip as determined experimentally at low speed for various spanwise and vertical locations and angles of attack as variations with chordwise distance. Results regarding the swept-wing-fuselage flow fields, unswept-wing-fuselage flow fields, and a comparison of wing-fuselage and fuselage-alone flow fields are provided.

Internal performance of normal-shock rectangular, circular, and scoop inlets and of external-compression inlets experimentally obtained with varying immersion in a turbulent boundary layer. Recoveries varied from about 95 percent of theoretical in the free stream to 80 percent with complete immersion, while the corresponding mass flows were usually above 95 percent of theoretical. Turning of the flow through 10 degrees caused losses in pressure recovery of 0.03 to 0.07. External compression did not improve pressure recovery in the boundary layer. Average distortion at critical operation for all inlets was 5 percent.

Memorandum presenting an investigation to determine the drag, longitudinal stability, and lateral stability characteristics of a model of a fighter-type airplane. Several modifications were made to the model in an attempt to eliminate pitch-up.

Report presenting the basic performance characteristics of several types of designs of subsonic-diffuser-bypass-duct combinations for bypass flows up to one-third of the total flow. The tested bypass-duct design procedure was found to be greatly oversimplified and was unable to produce satisfactory performance because of the extreme adverse pressure gradient.

A modified corrugated-insert blade with integrally cast shell and base was developed. This blade was as light as a conventional fabricated corrugated-insert blade. Of four test blades operated in a full-scale turbojet engine, one failed after about 15 hours operation at an inlet gas temperature of 1670 degrees F, a coolant-flow ratio of 0.0064, and a 1/3-span centrifugal stress of approximately 28,000 psi. Three other test blades ran for approximately 16, 31, and 36 hours without failure at similar conditions.

"The performance of a 2.8 KS 8100 Cajun solid-propellant rocket motor has been determined from the free-flight test of a single-stage model which reached a velocity of 5,268 feet per second (Mach number, 4.74). Thrust data from the flight test are compared with data obtained from a ground firing test made by the manufacturer and the two are shown to be in agreement." Report presenting general drag, data, and acceleration are also provided.

Memorandum presenting an investigation in the supersonic pressure tunnel at a free-stream Mach number of 2.01 and at angles of attack up to 15 degrees of the static pressures and boundary-layer characteristics on the forward parts of nine fusleages of various cross-sectional shapes. Results regarding the static pressure distributions, boundary-layer characteristics, and selection of fuselage shapes for use with top-mounted scoop inlets are provided.

Report presenting testing of the performance of a twin-duct air-intake system mounted on the sides of a 1/8-scale fuselage forebody model of a proposed aircraft at several Mach numbers, angles of attack, and yaw. Results regarding the inlet survey, compression-surface boundary-layer removal and effects of second-ramp position, instability, effect of cant on inlet performance, inlet performance with 0 and 30 degree second ramp angles, performance with fixed second-ramp angles, and effect of canards are provided.

Report presenting testing of a twin-duct air-intake system mounted on the sides of a fuselage forebody model at free stream Mach numbers of 1.5 and 2.0 over angles of attack and yaw. Results regarding the inlet survey, compression-surface boundary-layer removal and effects of second-ramp position, instability, effect of cant on inlet performance, inlet performance with 0 degree and 30 degree second-ramp angles, performance with fixed second-ramp angles, and effect of canards are provided.

"Drag characteristics of a series of related conical and circular-arc afterbodies are presented for Mach numbers from 0.6 to 1.3. Drag was obtained from pressure measurements on the boattail and solid base. The boattail angles tested ranged from 0 degrees to 45 degrees for ratios of base diameter to maximum body diameter ranging from 0 to 1.0" (p. 1).

Tests were performed in the Langley Unitary Plan wind tunnel to determine the drag and static longitudinal and lateral stability and control characteristics of a 1/20-scale model of the McDonnell F4H-1 airplane at Mach numbers of 1 57, 1 87, 2.16, and 2.53. This is the second phase in a series of tests performed on this model. The Reynolds numbers for these tests, based on the mean aerodynamic chord of the wing, are 1.446 x 10 (exp 6), 1.269 x 10 (exp 6), 1.116 x 10 (exp 6), and 0.714 x 10 (exp 6) at Mach numbers of 1.57, 1.87, 2.16, and 2.53, respectively.

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 investigation at transonic speeds to determine the effects of a sonic propulsive jet on the aerodynamic characteristics of the body from which it issues. A variety of afterbody shapes and their pressure distributions are described. Correlations are also drawn between certain characteristics of the afterbody and the effect of temperatures.

One-dimensional mean-section flow and blade specific losses proportional to average specific kinetic energy are assumed in the analysis. Range of the work-speed parameter lambda considered includes low to moderate blade speeds with high specific work outputs, where critical turbojet, turbopump, and accessory-drive turbines are encountered. A diffusion factor of 0.5 limits the loading on the downstream stators. Turbine efficiences considered are total or aerodynamic, rating, and static. Efficiences of velocity-diagram types at impulse and that corresponding to values of maximum efficiency are presented and compared to indicate in what range of lambda downstream stators are beneficial as well as the attending improvements in efficiency.

Report presenting an experimental investigation on a single turbojet engine combustor and on a turbojet engine using a fuel mixture of 70 percent trimethyl borate and 30 percent methyl alcohol. Results regarding the single combustor's deposition, combustion efficiency, ignition, temperature distribution, pressure drop and the engine operation are provided.

Report presenting static tests on the compatibility of gaseous fluorine with various liquids, solid plastics, waxes, and greases at pressures of 0 and 1500 pounds per square inch gage and atmospheric temperature. Several materials were found to be compatible at atmospheric pressure, but only Teflon and ruby were compatible under static conditions at 1500 pounds per square inch gage.

Static tests on the compatibility of fluorine with non-metals at atmospheric temperature eliminated many materials from further consideration for use in fluorine systems. Several materials were found compatible at atmospheric pressures. Only Teflon and ruby (aluminum oxide) were compatible at 1500 pounds per square inch gage.

Memorandum presenting an analysis of the effect of adding fluorine to the Vanguard first-stage oxidant. An increase in specific impulse of 5.74 percent may be obtained with 30 percent fluorine. Charts are provided regarding the vehicle performance increases for a zero-drag vertical trajectory.

From Introduction: "This report presents data pertinent to the problem of boosting rocket performance by adding up to 30 percent liquid fluorine to the liquid oxygen of an existing oxygen-hydrocarbon rocket engine."

Increasing the volume and length of the duct behind the inlet affected the inlet stability at Mach 2.0 and zero angle of attack. Close approximation of the inlet stability limit of the J34 engine-inlet configuration was obtained by a cold-pipe configuration having a length and volume approaching that measured to the engine turbine. Variation of these parameters had a small effect on the minimum subcritical stable mass flow below a cowl-lip-position parameter of 44 degrees and appeared to have a negligible effect on the inlet pressure-recovery - mass-flow curve. Initial buzz frequency and minimum cowl-lip-position parameter for complete buzz-free operation varied with configuration.

At a corrected speed of 1100 feet per second, the low-blade-angle rotor operated with a relative inlet Mach number of 1.2, a diffusion factor of 0.65, and an axial velocity ratio of 0.71 in the tip region (11 percent of passage height away from the outer wall). The measured minimum-loss coefficient was 0.35, and this value falls above a previous correlation of rotor losses with diffusion factor. Through a comparison with data for three other rotors, the occurrence of high losses was related to a high suction-surface Mach number. These comparisons also indicated that axial velocity ratios between 0.73 and 1.10 have no independent effect on losses.

Memorandum presenting an investigation with a ramjet utilizing pentaborane fuel in order to evaluate promising high-energy fuels under actual flight conditions. During descent, the engine accelerated to a Mach number of 1.45 when a flame-out believed due to spray-bar failure took place. Results regarding time history, pressure recovery, thrust, thrust-minus-drag, and drag coefficient, and comparison with hydrocarbon fuel are provided.

The high-work-output turbine had an experimental efficiency of 0.830 at the design point and a maximum efficiency of 0.857. The downstream stator was effective in providing axial flow out of the turbine for almost the whole range of turbine operation.