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NACA Conference on Aerodynamics of High Speed Aircraft

Description: This document contains reproductions of technical papers presented by staff members of the NACA Laboratories at the NACA Conference on Aerodynamics of High-Speed Aircraft held at the Langley Aeronautical Laboratory November 1, 2, and 3, 1955. The primary purpose of the conference was to convey to contractors of the military services and others concerned with the design of aircraft the results of recent research and to provide those attending with an opportunity to discuss these results. The papers in this document are in the same form in which they were orally presented at the conference to facilitate their prompt distribution. The original presentation and this record are considered as complementary to, rather than as substitutes for, the Committee's more complete and formal reports.
Date: November 1, 1955
Partner: UNT Libraries Government Documents Department

Analysis of a Pneumatic Probe for Measuring Exhaust-Gas Temperatures with Some Preliminary Experimental Results

Description: 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.
Date: May 21, 1952
Creator: Scadron, Marvin D.
Partner: UNT Libraries Government Documents Department

Investigation of Ejection Releases of an MB-1 Rocket from a 0.04956-Scaled Model of the Convair F-106A Airplane at Several Mach Numbers and Simulated Altitudes

Description: As a continuation of an investigation of the ejection release characteristics of an internally carried MB-1 rocket in the Convair F-106A airplane, fin modifications at additional Mach numbers and simulated altitudes have been studied in the 27- by 27-inch preflight jet of the Langley Pilotless Aircraft Research Station at Wallops Island, Va. The MB-1 rocket was ejected with fins open, fins closed, fins closed with a shroud around the fins, and fins folded with a "boattail" placed in between the fins. Dynamically scaled models (0.0^956 scale) were tested at simulated altitudes of 12,000, 18,850, and 27,500 feet at subsonic Mach numbers and at 18,850, 27,500, and 40,000 feet for Mach numbers of 1-39, 1-59, and 1.98. Successful ejections can be obtained for over 10 store diameters from release point by the use of a shroud around the folded fins with the proper ejection velocity and nose-down pitching moment at release. In one case investigated it was found desirable to close off the front one-third of the bomb bay. It appeared that the fins should be opened after release and within 5 "to 6 rocket diameters if no modifications are made on the rocket. An increase in fuselage angle of attack caused higher nose-up pitch rates after release.
Date: August 22, 1957
Creator: Lee, J. B. & Basford, R. C.
Partner: UNT Libraries Government Documents Department

Investigations of Air-Cooled Turbine Rotors for Turbojet Engines. 1: Experimental Disk Temperature Distribution in Modified J33 Split-Disk Rotor at Speeds up to 6000 RPM

Description: An experimental investigation is being conducted at the Lewis laboratory to establish general principles for the design of noncritical turbine rotor configurations. This investigation includes evaluation of cooling effectiveness, structural stability, cooling-air flow distribution characteristics, and methods of supplying cooling air to the turbine rotor blades. Prior to design of a noncritical rotor, a standard turbine rotor of a commerical turbojet engine was split in the plane of rotation and machined to provide a passage for distributing cooling air to the base of each blade. The rotor was fitted with nontwisted, hollow, aircooled blades containing nine tubes in the coolant passage. In the investigation reported herein, the modified turbine rotor operated successfully up to speeds of 6000 rpm with ratios of cooling-air to combustion-gas flow as low as 0.02. The disk temperatures observed at these conditions were below 450 0 F when cooling air at 100 F was used from the laboratory air system. The calculated disk temperatures based on the correlation method presented for rated engine conditions were well below 1000 F at a cooling-air flow ratio of 0.02, which is considered adequate for a noncritical rotor. An appreciable difference in temperature level existed between the forward and rear disks. This temperature difference probably introduced undesirable disk stress distributions as a result of the relative elongations of the two disks. This investigation was terminated at 6000 rpm so that slight changes in the engine configuration could be made to relieve this condition.
Date: January 9, 1952
Creator: Schramm, Wilson B. & Ziemer, Robert R.
Partner: UNT Libraries Government Documents Department

Experimental and Analytical Study of Balanced-Diaphragm Fuel Distributors for Gas-Turbine Engines

Description: A method of distributing fuel equally to a plurality of spray nozzles in a gas-turbine engine by means of balanced-diaphragm fuel distributors is presented. The experimental performance of three of eight possible distributor arrangements are discussed. An analysis of all eight arrangements is included. Criterions are given for choosing a fuel-distributor arrangement to meet specific fuel-system requirements of fuel-distribution accuracy, spray-nozzle pressure variations, and fuel-system pressures. Data obtained with a model of one distributor arrangement indicated a maximum deviation from perfect distribution of 3.3 percent for a 44 to 1 range (19.5 to 862 lb/hr) of fuel-flow rates. The maximum distributor pressure drop was 125 pounds per square inch. The method used to obtain the required wide range of flow control in the distributor valves consisted in varying the length of a constant-area flow path.
Date: August 14, 1950
Creator: Straight, David M. & Gold, Harold
Partner: UNT Libraries Government Documents Department


Description: The Neptune device was detonated underground in a room approximately 12 x 17 x 10 ft, at the end of a hooked drift. The yield was 115 plus or minus 15 tons. The shot and its effects are described and the major contributions of the data to the theory and prediction of cratering phenomenology are indicated. (W.D.M.)
Date: April 19, 1960
Creator: Shelton, A V; Nordyke, M D & Goeckermann, R H
Partner: UNT Libraries Government Documents Department

Investigation of the Pressure-Loss Characteristics of the Westinghouse X24-C-2 Inlet Screen, TED No. NACA 0447

Description: At the request of the Bureau of Aeronautics, Navy Department, investigations of the static-pressure losses and total-head distributions of the Westinghouse X24-C-2 inlet screen were made in the induction aerodynamics laboratory at Langley. The screen was investigated in two configurations, both before and after rounding the leading edges of the vanes. Investigations were conducted through air flows up to about pounds per second. The results of the investigations indicate that maximum lift coefficients of 1.36, 1.71 and 2.11 were measured on the model with flaps neutral and deflected 20 deg. and 55 deg, respectively, at a reynolds number of 8,600,000. When the duct inlet was replaced by a basic airfoil nose the flap neutral maximum lift coefficient was increased from 1.36 to 1.41. The results also showed that at maximum lift with flaps neutral or deflected 55 deg most of the area between the nacelles was stalled while only small areas on other portions of the model were stalled; when the duct inlet was replaced by the basic airfoil nose the stall was delayed to a slightly higher angle of attack but the nature of the stall was relatively unaffected.
Date: December 2, 1946
Creator: Lankford, John L.
Partner: UNT Libraries Government Documents Department

Investigation of an Experimental Supersonic Axial-Flow Compressor

Description: An investigation is in progress at the Langley Laboratory of the NACA to explore the possibilities of axial-flow compressors operating with supersonic velocities relative to the blade rows. The first phase of this investigation, a study of supersonic diffusers, has been reported. The second phase, an analysis of supersonic compressors, has also been reported. Preliminary calculations have shown that very high pressure ratios across a stage, together with somewhat increased mass flows, are possible with compressors which decelerate air through the speed of sound in their rotor blading. These performance characteristics are desirable in compressors for aircraft jet propulsion units, gas turbines, or superchargers. The third phase, presented here, is a preliminary experimental investigation of a supersonic compressor designed to produce a high pressure ratio in a single stage.
Date: August 7, 1947
Creator: Erwin, John R.; Wright, Linwood C. & Kantrowitz, Arthur
Partner: UNT Libraries Government Documents Department