Memorandum presenting an investigation of the effect of base bleed on the base force of a supersonic-interceptor model in the 8- by 6-foot supersonic wind tunnel. Generally, base pressures greater than ambient were obtained at the engine operating pressure ratios of the primary nozzle.

Sea-level static testing of turbojet engine afterburner.

From Summary: "This report presents the aerodynamic measurements made during a Navy investigation conducted in the Langley 8-foot high speed tunnel to determine the actual human tolerance to wind blasts."

Report presenting a comparison of the vertical-tail loads determined from pressure-distribution measurements in flight in various maneuvers with the corresponding vertical-tail loads. Some of the maneuvers investigated included slow rolls, steady sideslips, fishtails, and rolling pull-outs.

From Introduction: "Evaluation of tail-pipe burning in this engine with a larger tail-pipe combustion chamber is discussed in reference 1. Results of investigations on tail-pipe burning in this engine at static sea-level conditions are presented in reference 2. An investigation of thrust augmentation by means of injecting water at the inlet of an axial-flow compressor engine is discussed in reference 3."

Report presenting a flight test of a supersonic research pilotless aircraft in which large-amplitude aileron oscillations, most likely aileron compressibility flutter, were encountered in the transonic and supersonic speed ranges. Results regarding power-on flight and coasting flight are provided.

Report presenting an investigation in the propeller-research tunnel to develop wing-leading-edge inlets for locations between the inboard and outboard nacelles on each wing of a four-engine airplane. Testing was performed on the basic wing and original inlet as well as NACA-developed inlets for two versions of the airplane.

"An analysis was made at flight Mach number of 0.9 to estimate performance of nuclear-energy-powered turbojet engine and optimum engine operating conditions and to determine gross weight and load-carrying capacity of airplane powered by such an engine. The size of airplane required to carry disposable load of 20,000 pounds was found to vary from approximately 300,000 to 900,000 pounds depending on assumptions. For a reactor tube-wall mean temperature of 2500 degrees R, turbine-inlet temperature of 2000 degrees R, reactor-free-flow-area ratio of 0.33, reactor-shielding-material specific gravity of 6.0, shielding thickness of 3.0 feet, and altitude of 30,000 feet, the airplane gross weight required to carry a 20,000 payload is 545,000 pounds" (p. 1).

From Introduction: "Data are presented to show the effects of tail-pipe fuel-air ratio, altitude, and flight Mach number on tail-pipe-burner performance at rated engine speed and approximately constant turbine-outlet temperature. Operational characteristics of the tail-pipe burner and variable-area exhaust nozzle are also reported."

An equation for average heat transfer of a surface was derived when the boundary layer changed from laminar to turbulent. Influences on the heat transfer through a laminar boundary layer of Mach number, temperature ratio (gas temperature divided by wall temperature), and exponents of gas-property temperature relations were shown to be relatively small for air with Mach numbers less than 2 and temperature ratios between 1 and 4. Good agreement was obtained with experimental results from cylinders, an airfoil, and turbine blades.

Report presenting the results of a test program in a supersonic tunnel to determine the maximum lift of wings operating at supersonic speeds. A variety of wing plan forms of several thickness distributions were tested at a range of Mach numbers, Reynolds numbers, and angles of attack. The lift results, drag results, and Schileren photographs are described.

From Introduction: "The results of two phases of this investigation, presented in references 1 and 2, have indicated possible ways of reducing hull drag without causing large changes in aerodynamic stability and hydrodynamic performance."

Memorandum presenting an investigation of an NACA submerged intake installation on a modified fighter airplane conducted to determine the full-scale aerodynamic characteristics of this installation. Additionally, tests were conducted on the submerged inlet with revised entrance lips and deflectors to determine the configuration which would result in the best dynamic pressure recovery measured at the inlet for this installation without a major rework of the entrance.

Report presenting an investigation of the maximum hovering time, or the time that a helicopter can sustain itself without motion, for helicopters using a reciprocating engine, by ramjet engines at the tips of rotor blades, and by pulse-jet engines at the tips of rotor blades. Testing showed that the reciprocating engine permitted much longer hovering time than the jet-propulsion engines, but the jet-propulsion engines were much lighter and could lift greater disposable loads.

Report presenting an investigation of a series of flame holders designed with the objective of providing a high combustion efficiency in a ramjet combustor at an inlet air velocity of 200 feet per second, inlet-air pressure of 60 inches of mercury absolute, inlet air temperature of 200 degrees Fahrenheit, and near stoichiometric fuel-air ratio. Results regarding the preliminary experiments, two parallel rows of gutters, three rows of gutters, staggered gutters, simultaneous use of three rows and staggered gutters, use of molybdenum, and magnitude of pressure fluctuations are provided.

Report presenting a wind-tunnel investigation of a semispan model of a wing swept back 63 degrees with an aspect ratio of 3.5 and a taper ratio of 0.25. The tests were conducted to evaluate the effects of Reynolds and Mach number on the aerodynamic characteristics of the wing. Results regarding the characteristics of the wing with the elevon undeflected, effectiveness of the elevon, effects of roughness strips, and effect of model deflection under varying loads are provided.

Report presenting an investigation of several aileron modifications in conjunction with a tapered, sweptback wing with circular-arc airfoil sections of relatively large thickness ratio. This testing permits the evaluation of the wing-aileron rolling effectiveness over a range of Mach numbers. Results regarding true-contour ailerons, extended-chord ailerons, and blunt trailing-edge ailerons are provided.

Report presenting measurements of downwash angles at points at two vertical positions at the probable tail location behind a high-aspect-ratio wing with an NACA 65-210 section with no sweep and 30 degrees and 45 degrees of sweepback and sweepforward in conjunction with a fuselage.

Memorandum presenting an investigation to determine the automatic lateral stability characteristics of a model equipped with a gyro stabilizing unit that gave response to bank and yaw. Stable flights were obtained with the flicker-type automatic control, and the amplitude of the oscillations was decreased by adding the attachment which provided hunting control.

Report presenting an investigation to determine the aerodynamic characteristics of a wing with a leading sweptback edge at 47.5 degrees and a 20-percent-chord, 50-percent-span outboard aileron. The wing had symmetrical circular-arc airfoil sections and was tested with a circular-arc contour aileron and a flat-sided contour aileron with finite trailing-edge thickness. The longitudinal aerodynamic characteristics and aileron effectiveness are provided for both types of ailerons.

"A two-dimensional wedge diffuser, designed to be used with a ducted-airfoil ram jet, was tested on a rocket-powered test vehicle up to a Mach number of 1.45. Orifice plates and a choking section behind the diffuser exit simulated burning in a combustion chamber by providing the proper pressure drop. Results obtained from these tests showed there was no abrupt changes in mass flow and pressure recovery as the model velocity increased through the transonic region" (p. 1).

Performance characteristics of a two-dimensional isentropic diffuser have been experimentally determined at a Mach number of 3.85. At zero angle of attack, a maximum total-pressure recovery of 0.41 was obtained with a supercritical mass-flow ratio of 0.95. As a consequence of the twin-duct arrangement of the diffuser, a large discontinuity in pressure recovery and mass flow with a characteristic hysteresis was encountered between critical and subcritical operation. An asymmetric shock pattern with large-scale separation and flow reversal in one of the passages occurred at reduced mass flows. Pressure and force data presented for an angle-of-attack range from zero to 4 degrees.

Report presenting an investigation conducted to determine the overall performance of the J73 two-stage turbine with a standard rotor-blade configuration. The turbine operated with a maximum brake internal efficiency between 0.91 and 0.92 at an overall pressure ratio of about 3.4 and 120 percent equivalent design rotor speed. Results regarding the overall performance, variation of equivalent weight flow with overall pressure ratio, and effect of limiting blade loading are provided.

From Introduction: "The investigation reported herein is a continuation of a ram-jet-combustor design program being conducted at the NACA Lewis laboratory. The purpose of this broad program is to establish basic design criteria for combustors operating over wide range of fuel-air ratio with low pressure losses and high combustion efficiency, and to utilize these design criteria in the development of practical ram-jet combustors."