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 Collection: National Advisory Committee for Aeronautics Collection
Compatibility of Metals with Liquid Fluorine at High Pressures and Flow Velocities
No abstract available.
Comparison of Heat Transfer from Airfoil in Natural and Simulated Icing Conditions
An investigation of the heat transfer from an airfoil in clear air and in simulated icing conditions was conducted in the NACA Lewis 6- by 9-foot icing-research tunnel in order to determine the validity of heat-transfer data as obtained in the tunnel. This investiation was made on the same model NACA 65,2-016 airfoil section used in a previous flight study, under similar heating, icing, and operating conditions. The effect of tunnel turbulence, in clear air and in icingwas indicated by the forward movement of transition from laminar to turbulent heat transfer. An analysis of the flight results showed the convective heat transfer in icing to be considerably different from that measured in clear air and. only slightly different from that obtained in the icing-research tunnel during simulated icing.
Wind-tunnel investigation of several factors affecting the performance of a high-speed pursuit airplane with air-cooled radial engine
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Preliminary wind-tunnel investigation of the effect of area suction on the laminar boundary layer over an NACA 64A010 airfoil
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Preliminary Wind-Tunnel Investigation of the Performance of Republic F-105 Wing-Root Inlet Configurations at Various Angles of Attack and a Mach Number of 2.01
A 1/13-scale model of the forebody of the Republic F-105 with twin-duct wing-root inlets was tested in the Langley 4- by 4-foot supersonic pressure tunnel through a range of angle of attack from -4 deg to 15 deg at a Mach number of 2.01 and a Reynolds number of approximately 3.4 x 10(exp 6) per foot. The tests were made with four configurations which incorporated varying amounts of sweep and stagger of the inlet leading edges, modifications to the areas of the boundary-layer diverter floor plate, and modifications to the area of the boundary-layer diverter bleed slots. The highest overall pressure recovery at an angle of attack of 0 deg (average total-pressure recovery, 0.84 mass-flow ratio, 0.98) was achieved with configuration having an inlet leading-edge sweep angle of 58 deg with no leading-edge stagger. Stagger was found to improve the angle-of- attack performance, but at a sacrifice in inlet efficiency for an angle of attack of 0 deg. The boundary-layer diverter floor height, of the order of one boundary-layer thickness, was satisfactory for bypassing the fuselage boundary layer. The boundary-layer diverter-plate bleed slots were effective in increasing the total-pressure recovery of the inlet. The total-pressure-recovery contour plots, taken at the compressor-face station, indicate the existence of high-velocity "cores" throughout the inlet operating range.
Analysis of a Pneumatic Probe for Measuring Exhaust-Gas Temperatures with Some Preliminary Experimental Results
A pneumatic probe based on continuity of mass flow through two restrictions separated by a cooling chamber was constructed to measure gas temperature at and beyond the limit of thermocouples. This probe consisted of a subsonic flat-plate orifice for the first restriction and a sonic-flow converging-diverging nozzle for the second restriction. The effect of variation in gas constants on the calibration is examined for common engine-exhaust gases. A high-temperature wind tunnel that allowed calibration of the probe at temperatures up to 2000 deg R and. Mach numbers up to 0.8 is described. Agreement to better than 30 deg R between pneumatic probe indication and the indication of a rake of radiation shielded thermocouples indicates that extrapolation of the calibration to higher temperatures is possible with fair accuracy.
Adaptation of a Cascade Impactor to Flight Measurement of Droplet Size in Clouds
A cascade impactor, an instrument for obtaining: the size distribution of droplets borne in a low-velocity air stream, was adapted for flight cloud droplet-size studies. The air containing the droplets was slowed down from flight speed by a diffuser to the inlet-air velocity of the impactor. The droplets that enter the impactor impinge on four slides coated with magnesium oxide. Each slide catches a different size range. The relation between the size of droplet impressions and the droplet size was evaluated so that the droplet-size distributions may be found from these slides. The magnesium oxide coating provides a permanent record. of the droplet impression that is not affected by droplet evaporation after the. droplets have impinged.
Analogy Between Mass and Heat Transfer with Turbulent Flow
An analysis of combined heat and mass transfer from a flat plate has been made in terms of Prandtl t s simplified physical concept of the turbulent boundary layer. The results of the analysis show that for conditions of reasonably small heat and mass transfer, the ratio of the mass-and heat-transfer coefficients is dependent on the Reynolds number of the boundary layer, the Prandtl number of the medium of diffusion, and the Schmidt number of the diffusing fluid in the medium of diffusion. For the particular case of water evaporating into air, the ratio of mass-transfer coefficient to heat-transfer coefficient is found to be slightly greater than unity.
Analytical Investigation of Icing Limit for Diamond-Shaped Airfoil in Transonic and Supersonic Flow
Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.
The art of writing scientific reports
As the purpose of the report is to transmit as smoothly and as easily as possible, certain facts and ideas, to the average person likely to read it, it should be written in a full and simple enough manner to be comprehended by the least tutored, and still not be boring to the more learned readers.
Analysis of Meteorological Data Obtained During Flight in a Supercooled Stratiform Cloud of High Liquid-Water Content
Flight icing-rate data obtained in a dense and. abnormally deep supercooled stratiform cloud system indicated the existence of liquid-water contents generally exceeding values in amount and extent previously reported over the midwestern sections of the United States. Additional information obtained during descent through a part of the cloud system indicated liquid-water contents that significantly exceeded theoretical values, especially near the middle of the cloud layer.. The growth of cloud droplets to sizes that resulted in sedimentation from the upper portions of the cloud is considered to be a possible cause of the high water contents near the center of the cloud layer. Flight measurements of the vertical temperature distribution in the cloud layer indicated a rate of change of temperature with altitude exceeding that of the moist adiabatic lapse rate. This excessive rate of change is considered to have contributed to the severity of the condition.
The comparison of well-known and new wing sections tested in the variable density wind tunnel
Three groups of airfoils have been tested in the variable density wind tunnel. The first group contains three airfoils. The second group is a systematic series of twenty-seven airfoils. The third group consists of several frequently used wing sections.
Experimental investigation of effect of spike- tip and cowl-lip blunting on the internal performance of a two-cone cylindrical-cowl inlet at mach number 4.95
No Description
Experimental investigation of air-cooled turbine blades in turbojet engine. 2: Rotor blades with 15 fins in cooling-air passages
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Experimental Investigation of Effects of Combustion-chamber Length and Inlet Total Temperature, Total Pressure, and Velocity on Afterburner Performance
Afterburner combustion efficiency - effects of chamber length, temperature, pressure and velocity.
Experimental investigation of diffuser pressure ratio control with shock positioning limit on 28 inch ram-jet engine
No Description
Experimental performance of a 5000-pound- thrust rocket chamber using a 20-percent- fluorine - 80-percent-oxygen mixture with RP-1
Increased performance of rocket engine using fluorine-oxygen mixture with RP-1 fuel.
Experimental investigation of air-cooled turbine blades in turbojet engine. 1: Rotor blades with 10 tubes in cooling-air passages
No Description
Experimental investigation of extreme internal flow turning at the cowl lip of an axisymetric inlet at a Mach number of 2.95
No Description
Experimental Sea-level Static Investigation of a Short Afterburner
Sea-level static testing of turbojet engine afterburner.
Exploratory Investigation of Performance of Experimental Fuelrich Hydrogen Combustion System
Exploratory investigation of performance characteristics of fuel-rich hydrogen combustor.
Flight investigation of a liquid hydrogen fuel system
No Description
Performance of an Inlet Having a Variable-angle Two-dimensional Compression Surface and a Fixed-geometry Subsonic Diffuser for Application to Reduced Engine Rotative Speeds- Mach Numbers 0.66, 1.5, 1.7, and 2.0
Air inlet for turbojet engines having variable angle two-dimensional compression surface and fixed-geometry subsonic diffuser.
Performance of an All-internal Conical Compression Inlet With Annular Throat Bleed at Mach Number 5.0
Internal compression inlet with throat bleed-off at hypersonic flow.
Performance of an isentropic, all-internal- contraction, axisymmetric inlet designed for mach 2.50
Performance of internal contraction, axisymmetric inlet with isentropic compression surfaces on cowl and centerbody at Mach 2.0 to 2.7.
Performance of JP-4 fuel with fluorine- oxygen mixtures in 1000-pound-thrust rocket engines
Fluorine-liquid oxygen mixture and injector design effects on JP-4 jet fuel performance in rocket engines.
Performance of five short multielement turbojet combustors for hydrogen fuel in quarter-annulus duct
No Description
Performance of Isentropic Nose Inlets at Mach Number of 5.6
Isentropic nose inlet performance at mach 5.6.
Performance of several half-conical side inlets at supersonic and subsonic speeds
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Performance of twin-duct variable-geometry side inlets at mach numbers of 1.5 to 2.0
Supersonic wind tunnel test of twin-duct variable geometry side inlets.
Performance of Pentaborane, Pentaborane - Jp-4 Fuel Mixtures, and Trimethylborate Azeotrope Fuel in a Full-scale Turbojet Engine
Full scale turbojet engine testing of fuels - pentaborane-jp-4 mixtures and trimethylborate azeotrope performance.
Performance of Variable Two-dimensional Inlet Designed for Engine-inlet Matching. I - Performance at Design Mach Number of 3.07
Supersonic performance of variable two-dimensional inlets designed for engine-inlet matching.
High altitude performance investigation of J65-B-3 turbojet engine with both JP-4 and gaseous hydrogen fuels
No Description
High mach number, low-cowl-drag, external- compression inlet with subsonic dump diffuser
Low cowl drag, external compression inlet with subsonic dump diffuser for high Mach number application.
High-altitude performance of J71-A-11 turbojet engine and its components using JP-4 and gaseous-hydrogen fuels
No Description
Influence of boric oxide deposition on turbojet-engine operation
Influence of boric oxide deposition on turbojet engine operation.
Interference effects of fuselage-stored missiles on inlet duct model of an interceptor-type aircraft at Mach numbers 1.5 to 1.9
No Description
Interpretation of boundary-layer pressure-rake data in flow with a detached shock
No Description
Influence of Combustion-chamber Length on Afterburner Performance
Combustion-chamber length influence on turbojet engine afterburner performance.
De-icing Effectiveness of External Electric Heaters for Propeller Blades
No Description
Investigation at Mach number 1.91 of spreading characteristics of jet expanding from choked nozzles
No Description
Investigation of a 0.6 Hub-tip Radius-ratio Transonic Turbine Designed for Secondary-flow Study. Iii - Experimental Performance With Two Stator Configurations Designed to Eliminate Blade Wakes and Secondary-flow Effects and Conclusions From Entire Stator Investigation
Hub-tip radius ratio transonic turbine for secondary flow - stator configurations designed to eliminate blade wakes and secondary-flow effects.
Hydrogen for turbojet and ramjet powered flight
No Description
Measurement and analysis of turbulent flow containing periodic flow fluctuations
No Description
Low-temperature Chemical Starting of a 200-pound-thrust Jp-4 - Nitric Acid Rocket Engine Using a Three-fluid Propellant Valve
Low-temperature chemical starting of combined jp-4 nitric acid propellant for low-thrust rocket engine using three-fluid propellant valve.
Nonuniform Burnup and Poisoning Effects in a Reactor and Validity of Uniform Approximation
Burnup and poisoning in nuclear reactor core and validity of uniform approximation.
Measurement of distortion in second experimental control rod for argonne naval reactor with constant transverse temperature gradient and uniform longitudinal temperature distribution
No Description
Operation of an experimental air-cooled turbojet engine at turbine-inlet temperatures from 2200 R to 2935 R
No Description
Off-Design Performance of Divergent Ejectors
Off-design performance of divergent ejectors.
Performance and operational characteristics of pentaborane fuel in 48-inch-diameter ram-jet engine
Combustion efficiency of pentaborane fuel in 48- inch diameter ramjet engine.