National Advisory Committee for Aeronautics (NACA) - 46 Matching Results

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Location of Detached Shock Wave in Front of a Body Moving at Supersonic Speeds

Description: It is shown that for velocities slightly in excess of sonic, the position of detached shock wave located in front of a given body at zero angle of attack may be estimated theoretically to a reasonable degree of accuracy. In case of bodies of revolution the result was simple, but for two-dimensional bodies, pressure coefficient varies with Mach number, and slight difficulty appears. Theory developed compares favorably with available experimental data.
Date: May 6, 1947
Creator: Laitone, Edmund V. & Pardee, Otway O'm.
Partner: UNT Libraries Government Documents Department

Aerodynamic characteristics of a wing with quarter-chord line swept back 60 degrees, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section : transonic-bump method

Description: 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."
Date: September 6, 1949
Creator: King, Thomas J , Jr & Myers, Boyd C , II
Partner: UNT Libraries Government Documents Department

Aerodynamic characteristics at subcritical and supercritical Mach numbers of two airfoil sections having sharp leading edges and extreme rearward positions of maximum thickness

Description: 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."
Date: November 6, 1947
Creator: Eggers, A J , Jr
Partner: UNT Libraries Government Documents Department

The effect of blade-section thickness ratios on the aerodynamic characteristics of related full-scale propellers at Mach numbers up to 0.65

Description: Report discussing an investigation of two full-scale NACA propellers at a range of blade angles and at speeds of up to 500 miles per hour. The results are compared to previous investigations of five NACA propellers to evaluate the effects of blade-section thickness ratios on propeller characteristics.
Date: June 6, 1949
Creator: Maynard, Julian D. & Steinberg, Seymour
Partner: UNT Libraries Government Documents Department

Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine, 5, Combustion-Chamber Characterisitcs

Description: An investigation to determine the performance and operational characteristics of an axial-flow gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet ram-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and corrected horsepower. For the range of corrected engine speeds investigated, overall total-pressure-loss ratio, cycle efficiency, and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. For the range of corrected horsepowers investigated, the total-pressure-loss ratio and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horsepowers investigated at all corrected engine speeds.
Date: August 6, 1948
Creator: Geisenheyner, Robert M. & Berdysz, Joseph J.
Partner: UNT Libraries Government Documents Department

Tank Tests of an Alternate Hull Form for the Consolidated Vultee PB2Y-3 Airplane

Description: Tests have been made in Langley tank no. I of a dynamic model of the Consolidated Vultee PB2Y-3 airplane. These tests were made using an alternate hull form, the purpose of which was to reduce the bow spray and eliminate the landing instability which are objectionable features of the production design. The major differences from the PB2Y-3 hull included a deeper step to improve the landing stability , and a lengthened forebody and increased beam to reduce the sway in the propellers and on the flaps. The tests showed that the spray characteristics of the revised hull form were much better than that ot the production design. In addition the take-off and landing stability of the model with the alternate hull were satisfactory.
Date: November 6, 1946
Creator: Land, Norman S. & Posner, Jack
Partner: UNT Libraries Government Documents Department

Performance of J33 turbojet engine with shaft-power extraction III : turbine performance

Description: The performance of the turbine component of a J33 turbojet engine was determined over a range of turbine speeds from 8000 to 11,500 rpm.Turbine-inlet temperature was varied from the minimum required to drive the compressor to a maximum of approximately 2000 degrees R at each of several intermediate turbine speeds. Data are presented that show the horsepower developed by the turbine per pound of gas flow. The relation between turbine-inlet stagnation pressure, turbine-outlet stagnation pressure, and turbine-outlet static pressure was established. The turbine-weight-flow parameter varied from 39.2 to 43.6. The maximum turbine efficiency measured was 0.86 at a pressure ratio of 3.5 and a ratio of blade speed to theoretical nozzle velocity of 0.39. A generalized performance map of the turbine-horsepower parameter plotted against the turbine-speed parameter indicated that the best turbine efficiency is obtained when the turbine power is 10 percent greater than the compressor horsepower. The variation of efficiency with the ratio of blade speed to nozzle velocity indicated that the turbine operates at a speed above that for maximum efficiency when the engine is operated normally with the 19-inch-diameter jet nozzle.
Date: June 6, 1949
Creator: Huppert, M C & Nettles, J C
Partner: UNT Libraries Government Documents Department

Altitude-Test-Chamber Investigation of a Solar Afterburner on the 24C Engine I - Operational Characteristics and Altitude Limits

Description: 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. Over this range of operating conditions, the fuel-air ratio at which lean blow-out occurred was from 10 to 19 percent below these maximum fuel-air ratios. Combustion was very smooth and uniform during operation; however, ignition of the burner was very difficult throughout the investigation. A failure of the flame holder after 12 hours and 15 minutes of afterburner operation resulted in termination of the investigation.
Date: July 6, 1948
Partner: UNT Libraries Government Documents Department