"The preignition characteristics of a number of fuels have been studied under conditions similar to those encountered in an engine. These conditions were simulated by suddenly compressing a fuel-air mixture in contact with an electrically heated hot spot in the cylinder head of the NACA combustion apparatus. Schlieren photographs and indicator cards were taken of the burning, and the hot-spot temperatures necessary to cause ignition under various conditions were determined" (p. 121).

"Panels of aluminum alloy sheets, framed by side and end stiffeners, were subjected to combined loading by means of offset knife edges applying loads to top and bottom end plates with reacting forces against the end plates supplied by laterally acting rollers. The test specimens were 17S-T aluminum alloy shoots 0.040 inch thick in panels of 10-inch width and three different lengths (approximately 10, 26, and 30 inch). Data were obtained for the bowing of transverse and longitudinal ribs of rectangular cross section and varying depths mounted on one side of the sheet only, for several ratios of compression to shear loads" (p. 1).

"The results obtained from an evaluation for gust and draft velocities of acceleration and airspeed-altitude records taken by NACA recording instruments installed in P-61c airplanes participating in thunderstorm flights 6, 7, and 8 of July 9, 1946, July 10, 1946, and July 11, 1946, respectively, are presented herein. These data are summarized in tables I and II. In accordance with a recent discussion with a member of the U.S. Weather Bureau staff, the tabulated results for the present flight include in addition to data of the type presented in reference 1, the initial heading of the airplane for each traverse, the pressure altitude at the start of each traverse in increments of 500 feet, and the gust gradient distance when it could be evaluated" (p. 1).

Report presenting an investigation of the performance of a single can-type combustor designed for a turbojet engine equipped with an 11-stage axial-flow compressor and a single-stage turbine. The investigation was conducted to determine the altitude operational limits of the engine for two fuels, combustion efficiencies at various simulated conditions of altitude and engine speed, combustor-outlet temperature distribution for several altitudes at constant engine speed, and combustor total-pressure drop.

Report presenting an investigation to determine the control hinge moments and effectiveness at transonic speeds of a delta wing equipped with a constant-chord flap-type control with and without an unshielded triangular horn balance in the 7- by 10-foot tunnel using the transonic-bump method. The data indicated that the horn balance control was consistently more effective in changing lift at all Mach numbers than was the plain control but there was no appreciable difference in pitching-moment effectiveness.

Report presenting a fuselage and wing-fuselage combination with a wing of 45 degrees sweepback of the 0.25-chord line, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil sections in the slotted test section of a transonic tunnel for a range of Mach numbers and angles of attack. Results regarding lift characteristics, drag characteristics, lift-drag ratios, pitching-moment characteristics, a comparison with other test results, and base-pressure characteristics are provided.

Memorandum presenting an investigation at Mach numbers of 1.41, 1.62, and 1.96 to determine the effects of airfoil section and wing plan form on the pressures acting on the base of blunt-trailing-edge wings operating through an angle-of-attack range. The investigation included two groups of untapered wings of aspect ratio 2.7, with one group being unswept and the other with 45 degrees of sweepback. Results regarding the effects of wing location and body size, spanwise variations of base pressures, variations of average base pressure coefficients with angle of attack and wing section, and effects of plan form are provided.

Report presenting the lift of wing-body combinations at supersonic speeds as obtained by means of an approximate linearized-theory analysis involving an assumption concerning the geometry of the body. A comparison with several groups of data is provided to show the accuracy of the method.

"A preliminary investigation has been made in order to determine the static and dynamic longitudinal stability characteristics and the force characteristics of a Grunberg hydrofoil system comprising a main lifting hydrofoil and planing-surface stabilizers. The tests showed that a rearward movement of the center of gravity decreased the static and dynamic stability but increased the lift-drag ratio. In smooth water, adequate static stability was observed for all conditions tested" (p. 1).

Tests were made to determine the motor performance characteristics under flight conditions at propellant preignition temperatures of -25 degrees F, ambient temperature, and 130 degrees F. Results regarding the trajectories of the various test vehicles, time histories of velocity and Mach number, acceleration time histories, thrust curves, and maximum variation of total impulse are provided.

Report presenting hinge-moment data for a wing-body combination at Mach numbers of 1.46 and 1.99 and some lift and drag results for the lower Mach number. The model consisted of an aspect ratio 4, triangular, all-movable wing with an 8-percent-thick double-wedge section mounted on a pointed cylindrical body of fineness ratio 9-1/3. The results indicated that the wing has a relative effectiveness ranging from 63 to 98 percent.

The present status of available information relative to the prediction of shock-induced boundary-layer separation is discussed. Experimental results showing the effects of Reynolds number and Mach number on the separation of both laminar and turbulent boundary layer are given and compared with available methods for predicting separation. The flow phenomena associated with separation caused by forward-facing steps, wedges, and incident shock waves are discussed. Applications of the flat-plate data to problems of separation on spoilers, diffusers, and scoop inlets are indicated for turbulent boundary layers.

Combined compressor coolant injection and afterburning for turbojet engine thrust augmentation.

Report presenting a discussion of some of the requirements and of the application of the augmented performance of compressor coolant injection and afterburning in a turbojet engine in a high-speed, high-altitude flight. Results regarding the experimental performance of a combined system and the application of augmented performance to high-speed flight are provided.

Report presenting comparisons of low-lift drag measurements on eight airplane configurations at a range of Mach numbers by various techniques. Data was obtained from wind tunnel testing, rocket propelled model tests, and airplane flight tests. The agreement of data is generally good and no consistent Reynolds number effects are observed.

Report presenting comparisons of low-lift drag measurements on eight airplane configurations at a range of Mach numbers using various techniques. Data was obtained from wind-tunnel testing, rocket-propelled model testing, and airplane flight testing. The majority of the data showed agreement and discrepancies were likely due to issues regarding surface condition, individual test setup and accuracy, and detail geometric differences between the airplanes and respective models.

Report presenting an experimental vaporizing annular combustor that evolved from a high-performance annular combustor developed for vapor-fuel injection. The combustor consisted of a one-quarter sector of a single-annulus combustor designed to fit in a housing with an outside diameter of 25.5 inches, an inside diameter of 10.6 inches, and a combustor length of approximate 23 inches. Results regarding the performance of the Model 29I combustor, performance of Model 30, and performance of the best combustor configuration are provided.