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  • AERODYNAMICS  (4)
  • METEOROLOGY AND CLIMATOLOGY  (1)
  • 1975-1979  (5)
  • 1930-1934
  • 1
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    Supersonic inlet contour interpolation (1975)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: A method for designing supersonic inlet contours is described which consists in the interpolation of the contours of two known inlets designed for different Mach numbers, thereby determining the contours for a third inlet at an intermediate design Mach number. Several similar axisymmetric inlet contours were interpolated from known inlets with design Mach numbers ranging from 2.16 to 4.0 and with design Mach numbers differing by as much as 1.0. The flowfields were calculated according to Sorensen's (1965) computer program. Shockwave structure and pressure distribution characteristics are shown for the interpolated inlets. The validity of the interpolation is demonstrated by comparing the plots of the flowfield properties across the throat station of the interpolated inlet with the known inlets which were designed iteratively. It seems possible to write a computer program so that a matrix of known inlet contours can be interpolated.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 12; Sept
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Investigation of two bifurcated-duct inlet systems from Mach 0 to 2.0 over a wide range of angles of attack (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 15.354 percent/scale lightweight fighter type inlet/forebody was tested over a Mach number range of 0 to 2.0. Model configurations consisted of side mounted normal shock and fixed overhead ramp type inlets. Each configuration consisted of two inlets ducted (bifurcated) to supply a single engine face. The normal shock inlet variables included a boundary layer splitter bleed system, alternate boundary layer splitter plates, alternate upper and lower cowl lip shapes, and a blow-in-door (auxiliary inlet) in one lower lip. The only variable of the fixed overhead ramp inlet was the boundary layer bleed flow. Reynolds numbers ranged from 7.6 x 1 million to 19.5 x 1 million/m. Angle of attack ranged from -10 to 35 deg and angle of sideslip from -8 to 8 deg. Test measurements included engine face total pressure recovery, steady state distortion, dynamic distortion, and surface static pressures on the forebody and inlet surfaces.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-73118 , A-6512
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Field studies of the electrification of thunderstorms (1979)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Many theories have been advanced to explain the development of electric fields in thunderstorms, culminating in lightning, but thorough appraisal of these has been hampered by the lack of reliable and comprehensive observational data on the electrical characteristics, microphysical properties and dynamical behavior of the storms. A major field experiment (the Thunderstorm Research International Project) has been in progress for three years, in an effort to remedy this deficiency, and this paper describes some of this work and the results emanating from it. Major tools in this investigation are: an instrumented aircraft capable of penetrating the clouds; dual-Doppler and fast scanning radars; field-change and precipitation-recording networks; and an acoustic system for reconstructing the location of points on the lightning channels. The early results indicate a strong correlation between updraughts, precipitation and high fields. Circumstantial evidence points towards the presence of ice as being crucial to rapid field growth.
    Keywords: METEOROLOGY AND CLIMATOLOGY
    Type: Electrostatics 1979; Apr 17, 1979 - Apr 20, 1979; Oxford
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Transonic performance of an auxiliary airflow system for axisymmetric inlets (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A large-scale model of an axisymmetric inlet with a centerbody auxiliary airflow system has been tested in the wind tunnel at transonic speeds. The auxiliary system allows additional airflow (other than in the main duct formed by the cowl and translating centerbody) to pass through the centerbody of the inlet and combine with the main duct airflow on its way to the engine face. The results of the tests are presented, and the inlet performance is compared to a closely related alternative inlet with a 'traveling' boundary-layer bleed system which precludes the use of a centerbody auxiliary airflow system. The comparison shows that the auxiliary airflow inlet can supply 7.7% more engine face airflow at Mach number 1.0 and is 26% shorter than the traveling bleed inlet. Even though maximum transonic airflow was not achieved at a comparable engine face mass-flow ratio of 0.580, a total-pressure distortion of 0.10 and a total-pressure recovery of 0.985 were achieved for the auxiliary airflow inlet while a recovery of only 0.965 was achieved for the traveling bleed inlet.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 77-148 , Aerospace Sciences Meeting; Jan 24, 1977 - Jan 26, 1977; Los Angeles, CA
    Format: text
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    NASA TECHNICAL REPORTS
  • 5
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    Investigation of Inlet Concepts for Maneuver Improvement at Transonic Speeds (1977)
    In:  CASI
    Details
    Publication Date: 2019-08-14
    Description: A 15 percent scale lightweight fighter type inlet forebody was tested in the Ames 14 foot transonic wind tunnel at Mach numbers of 0.7, 0.9, and 1.04. The inlet was a two dimensional horizontal ramp system designed for a Mach number of 2.2. Four inlet devices designed to prevent or delay cowl-lip boundary layer separation or to improve the inlet internal flow characteristics at high angles of attack were investigated. The devices used to control cowl-lip separation consisted of cowl leading edge flaps, slotted flaps, and tangential blowing. To improve the internal flow characteristics, discrete jet nozzle flows were directed downstream and parallel to the duct surface in the subsonic diffuser to energize the wall boundary layer. The discrete jets used in the subsonic diffuser were also tested in combination with each of the cowl leading edge devices. Test measurements included engine-face total pressure recovery, steady state distortion, dynamic distortion, duct boundary layer profiles, and duct-surface static pressures.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-73215 , A-6952
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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