From Summary: "Tests of a 1/7 size model of the Grumman XJR2F-1 amphibian were made in Langley tank no.1 to examine the landing behavior in rough water and to measure the normal and angular accelerations experienced by the model during these landings. All landings were made normal to the direction of wave advance, a condition assumed to produce the greatest accelerations. Wave heights of 4.4 and 8.0 inches (2.5 and 4.7 ft, full size) were used in the tests and the wave lengths were varied between 10 and 50 feet (70 and 350 ft, full size)."

Report discussing the results of an investigation to determine the stability and control characteristics of a flying model of an unswept-wing vertically rising airplane. Information about stabilization, maneuverability, rolling motions, and landings is provided.

From Introduction: "A 12-percent-chord-thick wedge section and a reversed NACA 0012 section were chosen for these tests as they are representative of sections having no boat tailing and appreciable boat tailing (i.e., blunt and rounded trailing edges, respectively), and the results of this investigation are compared with those obtained from a previous investigation of the NACA 0012 section. Conclusions are drawn regarding the relative merits of the two unconventional sections and the conventional section in transonic speed range."

From Introduction: "Presented in this paper are the results of an investigation of a semispan model of a delta wing with 60^o sweepback at the leading edge which was equipped with a large triangular control having an overhang balance mounted on a skewed hinge axis. The purpose of this investigation was to determine the aerodynamic characteristics of a delta wing with a control which was designed to provide aerodynamic balance at zero control deflection based on the span load distribution of reference 1."

From Introduction: "The wings were tested in combination with fuselage similar to the one used in the 7- by 10-foot wind-tunnel investigations. The results are reported herein and are compared with results for three similar model wings on the transonic bump (references 1, 2, and 3)."

"A Douglas C-74 airplane, during a test dive at about 0.525 Mach number, experienced uncontrollable longitudinal oscillations sufficient to cause shedding of the outer wing panels and the subsequent crash of the airplane. Tests of a section of the horizontal tail plane from a C-74 airplane were conducted in the Ames 16-foot high-speed wind tunnel to investigate the possibility of the tail as a contributing factor to the accident. The results of the investigations of fabric-covered elevators in various conditions of surface deformation are presented in this report" (p. 1).

From Introduction: "This paper presents the results of the investigation of the wing-alone and wing-fuelage configurations employing a wing with the quarter-chord line swept back 60^o, aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil section parallel to the free stream."

Report discussing an investigation of the effect of sweepback angle on wing-fuselage characteristics at a subsonic and a supersonic Mach number. Lift, drag, and pitching-moment coefficients, downwash-angle, and total-pressure measurements are all provided.

From Introduction: "The heat-transfer data calculated from measured temperatures are compared with values calculated by the theory of Van Driest for a flat plate with laminar and turbulent boundary layers. In addition, the heat-transfer data from the flight tests are compared with data obtained from the Langley Structures Research Division of ground tests of an identical wing at a Mach number approximately equal to 1.99 in the pre-flight jet of the Langley Pilotless Aircraft Research Station at Wallops Island, Va. The stream static pressure is maintained at about 1 atmosphere, the free-stream temperature at about 75^o F, and the stagnation temperature at approximately 500 ^o F (ref.1)."

Testing was performed at Mach number 1.53 with a wing-fuselage combination with a wing with 63 degrees leading-edge sweep, an aspect ratio of 3.46, and a taper ratio of 0.25. The wing had an NACA 64A005 thickness distribution parallel to the plane of symmetry and was cambered and twisted. Results regarding the comparison of lift, drag, and pitching-moment characteristics of WF-63c and WF-63, effects of sweep, and effects of Reynolds number are provided.

Report presenting an experimental investigation to determine the aerodynamic characteristics of a slender cone-cylinder body of revolution in the 1- by 1-foot supersonic wind tunnel. Viscous drag and three component forces were measured at Mach number 3.12 and a range of Reynolds numbers and angles of attack. Results regarding pressure distributions, boundary layer, and forces are provided.

Pressure distributions for a series of four boattailed bodies of revolution were obtained and compared with theory for a Mach number of 3.12, a Reynolds number range of 2 x 10 to 6th power to 14 x 10 to the 6th power, and angles of attack from zero to 9 degrees. Second-order theory adequately predicted the pressure distribution for regions free of the effects of cross-flow separation.

Report presenting an investigation to determine the performance of 17 cooling-air ejectors at primary-jet pressure ratios from 1 to 10, secondary pressure ratios to 4.0, and a temperature ratio of unity. Results regarding pumping characteristics, effect of spacing ratio on pumping, effect of diameter ratio on pumping, thrust characteristics, effect of spacing ratio on thrust, effect of diameter ratio on thrust, comparison of cylindrical and conical ejectors, comparison of performance with small model ejectors, and a comparison of performance with full-scale installation are provided.

An investigation of the effect of inlet pressure, corrected engine speed, and turbine temperature level on turbine-inlet gas temperature distributions was conducted on a J40-WE-6, interim J40-WE-6, and prototype J40-WE-8 turbojet engine in the altitude wind tunnel at the NACA Lewis laboratory. The engines were investigated over a range of simulated pressure altitudes from 15,000 to 55,000 feet, flight Mach numbers from 0.12 to 0.64, and corrected engine speeds from 7198 to 8026 rpm.

From Introduction: "The results of this complete evaluation of the altitude performance and operational characteristics of the two types of flame-holder and fuel-system configurations are reported herein.The starting limits of both configurations at a flight Mach number of 0.6 are also discussed."

"Ignition of the fuel-air mixture in an afterburner of turbojet engine at altitude has often proved to be a different problem to solve. Electrical ignition has not proven satisfactory because of the unreliability of such systems (ref. 1). The ignition data reported herein were obtained for two after-burner configurations. Autoignition data are included for both configurations and hot-streak-ignition data, for only one. "

"An altitude-test-chamber investigation was conducted to determine the operational characteristics and altitude blow-out limits of a Solar afterburner in a 24C engine. At rated engine speed and maximum permissible turbine-discharge temperature, the altitude limit as determined by combustion blow-out occurred as a band of unstable operation of about 8000 feet altitude in width with maximum altitude limits from 32,000 feet at a Mach number of 0.3 to about 42,000 feet at a Mach number of 1.0. The maximum fuel-air ratio of the afterburner, as limited by maximum permissible turbine-discharge gas temperatures at rated engine speed, varied between 0.0295 and 0.0380 over a range of flight Mach numbers from 0.25 to 1.0 and at altitudes of 20,000 and 30,000 feet" (p. 1).

Altitude-test-chamber investigation of effects of flame-holder blocked area and gutter width on performance of 28-inch diameter ram jet at simulated flight Mach number of 2.0 for altitudes from 40,000 to 55,000 feet was conducted at NACA Lewis laboratory. Ten flame holders investigated covered gutter widths from 1.00 to 2.50 inches and blocked areas from 40.5 to 62.0 percent of combustion-chamber area. Gutter width did not appreciably affect combustion efficiency.

Report presenting a direct-connect altitude test chamber investigation of the combustion performance of a 28-inch-diameter ramjet engine with a can-type center pilot burner. Combustion-chamber configurations with three different flame holders were investigated at a simulated flight Mach number of 2.0 and several different altitudes.

From Summary: "An investigation was conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of an axial flow-type turbojet engine with a 4000-pound-thrust rating over a range of pressure altitudes from 5,000 to 50,000 feet, ram pressure ratios from 1.00 to 1.86, and temperatures from 60 deg to -50 deg F. The low-flow (standard) compressor with which the engine was originally equipped was replaced by a high-flow compressor for part of the investigation. The effects of altitude and airspeed on such operating characteristics as operating range, stability of combustion, acceleration, starting, operation of fuel-control systems, and bearing cooling were investigated."

From Introduction: "As part of a comprehensive investigation of the J40 turbojet engine conducted at the NACA Lewis altitude wind tunnel, the steady-state engine performance of the prototype J40-WE-8 turbojet engine without afterburner was obtained and is presented herein. A basic redesign of the compressor and other modifications in the compressor and the combustor were incorporated in the XJ40-WE-6 turbojet engine (references 2 and 3). In this report the modified engine is designated "the prototype J40-WE-8 without afterburner."

Report presenting analytical estimates of the thrust, efficiency, drag, and weight of the air-turborocket engine. The effects of changes in the engine components on design and off-design performance are emphasized. Results regarding the performance, effect of design parameters, effect of subsonic cruising Mach number, geometry variations, engine weight estimates, and comparison with turbojet engines are provided.

Report presenting airspeed and altitude data for nonscheduled cargo operations of a Douglas DC-4 airplane between Seattle, Washington and Alaska during February and March of 1947 to determine the probability of reaching or exceeding given values of airspeed and Mach number. The total probability of exceeding a placard never-exceed speed of 266 mph depends primarily on exceeding the speed in descent. The probability was found to be about once in 100 hours of descent or once in 1000 hours of general flying time.

From Introduction: "In the present paper, calculations are presented to show the relative effects of wide variations in the roll performance and normal-accleration capability. Brief analyses are also included to show the effect on the success of attacks of other design factors such as speed and radar range."