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  • 1
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    Prediction of supersonic and hypersonic inlet flow fields (1970)
    In:  CASI
    Details
    Publication Date: 2005-11-27
    Description: Computer programs for predicting flow field properties in supersonic and hypersonic inlets
    Keywords: AERODYNAMICS
    Type: L. METHODS IN AIRCRAFT AERODYN. 1970; P 583-595
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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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
  • 3
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    Advanced supersonic inlet technology. (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Recently, relatively new analytical procedures have been successfully used to design bleed systems for mixed-compression inlets designed to operate efficiently up to Mach number 2.65. The procedures used constitute a major advance in inlet technology by offering a promising approach to attain high internal and external performance for mixed-compression inlets that operate over a large supersonic Mach number range. Unfortunately, there is a lack of data describing bleed hole performance characteristics to verify these procedures at high Mach numbers. This paper briefly discusses the analytical procedures for designing advanced inlet systems and suggests facility modifications wherein the procedures can be verified on large-scale inlet models up to approximately Mach number 4.5.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 10; May 1973
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Boundary layers and hypersonic inlet flow fields. (1966)
    In:  Other Sources
    Details
    Publication Date: 2019-05-30
    Description: Shock wave boundary layer interactions and shear flow regimes in hypersonic inlet flows
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 66-606
    Format: text
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    NASA TECHNICAL REPORTS
  • 5
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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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  • 6
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    Investigation of a mixed compression axisymmetric inlet at Mach number 5.3 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The hypersonic diffuser portion of an uncooled high performance mixed compression, axisymmetric inlet suitable for subsonic burning engines was designed and tested. Performance of a model with a 25.4-cm capture diameter was measured in a wind tunnel and the results were compared with theoretical predictions calculated by a comprehensive computer program. All tests were conducted at a Mach number of 5.3 at a total temperature of 667 K and a total pressure of 11.57 atm. The angle of attack ranged from 0 to + or - 3 deg. Performance at angle of attack remained high. Reasonably high performance in the throat (maximum throat pitot-pressure recovery of 77 percent and an average value of 58 percent) was obtained at 0 deg angle of attack with relatively large amounts of boundary-layer bleed (11 to 22 percent of the capture mass flow). The computer program used in the design of this inlet is considered marginally adequate for predicting hypersonic inlet flow fields. Although the program as it now exists is very useful, an improved computer program that more accurately predicts the boundary layer and the shock-wave-boundary-layer interaction and accounts for boundary-layer bleed should be developed for reliability predicting hypersonic inlet flow fields.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6647 , A-4160
    Format: application/pdf
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  • 7
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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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  • 8
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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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