From Introduction: "The single purpose of this paper is to make available the data obtained from tests of these two Clark Y section propellers as quickly as possible with no attempt being made to analyze the results or to compare them with other high-speed-propeller test results."

From Introduction: "The present report gives all the results, including those published in reference 1 and the results of previously un-published tests tests of the airfoils at positive angles of attack."

An investigation has been conducted at the Langley 4- by 4-foot supersonic pressure tunnel at a Mach nmber of 2.01 to determine the aerodynamic characteristics of several configurations of a model of a 45 deg swept-wing airplane. The basic configuratin had a wing with 45 deg sweepback at the quarter-chord line, aspect ration 3.2, taper ration 0.468, NACA 65A005.5 sections just outboard of the inlet and NACA 65A003.7 sections at the tip. The wing was mounted slightly above the body center line and an all-movable horizantal tail was located slightly below the extended chord line of the wing. Tre design incorporated twin wing-root supersonic inlets ducted to a single exit at the base of the fuselage. The configurations investigated included an extended nose length, a bumped-fuselage afterbody, an inlet droop, an lncreased wing aspect ratio, and a revised canopy shape. Configurations employing the wing of increased aspect ratio of 3.7, which constituted the bulk of the tests, produced about a 10-percent increase in lift and in longitudinal stability as compared with the basic wing of aspect ratio 3.2. There was a slight but masurable increase in minimum drag and maximum lift-drag ratio.

Report presenting the aerodynamic heat-transfer properties and zero-lift drag of a typical pilot's canopy as determined by a rocket-model flight test through a range of Mach and Reynolds numbers. The canopy had a 63 degree sweptback flat windshield, circular cross section, and an equivalent body fineness ratio of 7.0. Results regarding the canopy temperature distribution, canopy pressure distribution, heat-transfer coefficient, and drag are provided.

Measured values of lift, drag, and pitching moment at a Mach number of 1.53 and Reynolds numbers of 0.31, 0.62, and 0.84 million are presented for a wing-fuselage combination having a wing leading-edge sweep angle of 63 degrees, an aspect ratio of 3.42, a taper ratio of 0.25, and an NACA 64A006 section in the stream direction. Data are also presented for sweep angles of 57.0 degrees, 60.4 degrees, 67.0 degrees, and 69.9 degrees. The experimentally determined characteristics were less favorable than indicated by the linear theory but the experimental and theoretical trends with sweep were in good agreement. Boundary-layer-flow tests showed that laminar boundary-layer separation was the primary cause of the differences between experiment and theory.

From Introduction: "Because of variation in the static-pressure drop across the porous material and in the cooling air required along the length of afterburner combustion chambers, a prescribed distribution of permeability is usually necessary. However, as a preliminary investigation, this report is concerned only with the attainment of uniform permeability. The results of an experimental investigation, conducted at the NACA Lewis laboratory, to determine the permeabilities of five thick meshes of wire filter cloth are presented herein."

Report presenting an altitude-chamber investigation to determine the altitude performance characteristics of the British Rolls-Royce Nene II turbojet engine with an 18.41-inch-diameter jet nozzles. Testing occurred at a range of simulated altitudes and ram-pressure ratios. Results regarding the simulated flight performance, generalized performance, and effect of jet-nozzle area on performance are provided.

From Introduction: "In response to the ever-increasing need for high thrust augmentation, an investigation was conducted that had as its primary objective the attainment of maximum exhaust-gas temperature and thrust (ref. 1). The investigation reported herein was therefore conducted to ascertain the operational limits of the most promising high-temperature afterburner design of reference 1 and to determine its performance over a wind range of flight conditions."

From Introduction: "During previous investigations of afterburning conducted at the NACA Lewis laboratory (references 1 to 4, for example), the objective of obtaining high exhaust-gas temperatures and, consequently, maximum thrust augmentation was compromised to some extent in order to alleviate the problem of afterburner shell cooling. The primary objective of the investigation reported herein was therefore the attainment of maximum exhaust-gas temperature imposed by the cooling techniques previously used, a secondary flow of cooling air was provided around the afterburner shell."

"An investigation was conducted in the Cleveland altitude wind tunnel to determine the performance of a Curtiss propeller with four 838-1C2-1SR1 blades on a YP-47M airplane at high blade loadings and engine powers. The study was made for a range of power coefficients between 0.30 and 1.00 at free-stream Mach numbers of 0.40 and 0.50. The results of the force measurements indicate primarily the trend of propeller efficiency for changes in power coefficient or advance-diameter ratio, inasmuch as corrections for the effects of tunnel-wall constriction on the installation have not been applied" (p. 1).

"An altitude-wind-tunnel investigation has been made to determine the performance of a Curtiss 732-1C2-0 four-blade propeller on a YP-47M airplane at high blade loadings and engine power. Propeller characteristics were obtained for a range of power coefficients from 0.30 to 1.00 at free-stream Mach numbers of 0.40 and .50" (p. 1).

From Summary: "A blade-element theory for axial-flow compressors has been developed and applied to the analysis of the effects of basic design variables such as Mach number, blade loading, and velocity distribution on compressor performance. The relations among several efficiencies useful in compressor design are derived and discussed. The possible gains in useful operating range obtainable by the use of adjustable stator blades are discussed and a rapid approximate method of calculating blade-angle resettings is shown by an example."

From Introduction: "Wind-tunnel and flight tests have shown that conical leading-edge camber on a thin low-aspect-ratio delta wing results in increasing the lift-drag ratio at transonic and low supersonic speeds (refs. 1 and 2). References 3 and 4 present the results of two previous investigations of this general program. A more detailed analysis of the pressure distributions of reference 5 is presented herein in terms of total section loads and overall wing-body characteristics."

Report presenting an examination of the limitations of the use of jet momentum coefficient as a correlating factor in comparing tests of blowing-type boundary-layer control. Theoretical and experimental data indicate that it should be an acceptable parameter where the duct pressures are large.

An analytical investigation was made of two supersonic interceptor type airplanes to determine the most desirable turbojet engine characteristics for this application The airplanes were designed differently primarily because of the amount of subsonic flight incorporated in the flight plan--one flight having none and the other, a cruise radius of 400 nautical miles. Several power plant design variables were varied independently to determine the effect of changes in each parameter on airplane performance. These parameters included compressor pressure ratio, compressor efficiency, turbine-inlet temperature, afterburner temperature, engine specific weight, and air-handling capacity. The effects of using a convergent-divergent exhaust nozzle and of changing the design flight Mach number were also investigated.

Memorandum presenting an investigation of coolant-flow rates for a turbine rotor blade with return-flow type of coolant-passage configuration formed by fins within a capped blade shell with both hydrogen and methane fuels as coolants. Results regarding spanwise blade and coolant temperature distributions, effects of coolant inlet temperature on hydrogen-coolant-flow requirements, comparison of return-flow-blade coolant requirements, effect of fin thickness and fin height, comparison of hydrogen and methane as coolants, and feasibility of fuel-cooled turbines are provided.

From Introduction: "The purpose of this report was to investigate the possible use of engine fuels (hydrogen and methane) as coolants for turbine rotor blades and to determine the pressure-drop characteristics of these coolants for a turbine blade with a more effective coolant-passage configuration than that considered in reference 5 but for the same engine and flight conditions."

From Introduction: "The results of tests in which area suction was applied near the leading edge of the wing or the knee of a leading-edge flap are reported in references 1 through 5."

"A theoretical analysis is presented for obtaining by use of Theodorsen's propeller theory the load distribution along a propeller radius to give the optimum propeller efficiency for any design condition. Examples are included to illustrate the method of obtaining the optimum load distributions for both single-rotating and dual-rotating propellers" (p. 1).

Report presenting an evaluation of a number of factors affecting engine-airframe arrangements in terms of range. Appropriate equations are developed and evaluated for a range of Mach numbers, ramjet and turbojet engines at several cycle temperatures, and two airplane lift-drag ratios. Some of the factors explored include inlet locations, jet cant for lift, engine moments for trim, a combination of the three factors, and the use of boundary layer in the engine.

Report presenting base pressures measured on several fin-stabilized bodies of parabolic-arc profile in free flight at a range of Mach and Reynolds numbers. The bodies varied in length but had the same afterbody ratios. Results regarding base pressure coefficients, side pressure coefficients, effect of afterbody length, and drag are provided.

Memorandum presenting an analysis of a system consisting of a mercury turbine-driven air compressor a mercury condenser wherein heat was added to the compressed air. The heat addition to the mercury is accomplished in an intermediate heat exchanger (mercury boiler) which has a liquid metal, other than mercury, circulating through the opposite side and through a nuclear reactor. Results regarding the turbine-exhaust pressure, turbine-inlet pressure, condenser-inlet Mach number, effect of nacelle drag, and some general data are provided.

Report presenting a new blade series designed to operate at near choking inlet Mach numbers and to have near maximum critical Mach numbers by employing high aerodynamic loading of the profiles in the leading-edge region and relatively straight trailing edges. Results regarding the aerodynamic characteristics and prediction of high-speed performance are provided.

Performance of experimental annular turbojet combustors at simulated high altitude supersonic flight conditions.