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  • 1
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    Evaluation of Flush-Mounted, S-Duct Inlets With Large Amounts of Boundary Layer Ingestion (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability, an experimental investigation of four S-duct inlet configurations with large amounts of boundary layer ingestion (nominal boundary layer thickness of about 40% of inlet height) was conducted at realistic operating conditions (high subsonic Mach numbers and full-scale Reynolds numbers). The objectives of this investigation were to 1) develop a new high Reynolds number, boundary-layer ingesting inlet test capability, 2) evaluate the performance of several boundary layer ingesting S-duct inlets, 3) provide a database for CFD tool validation, and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a fullscale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of this investigation indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion (by decreasing inlet throat height and increasing inlet throat width) or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise.
    Keywords: Aerodynamics
    Type: Paper 38 , Symposium on Vehicle Propulsion Integration; Oct 06, 2003 - Oct 09, 2003; Warsaw; Poland
    Format: application/pdf
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  • 2
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    Isolated Performance at Mach Numbers From 0.60 to 2.86 of Several Expendable Nozzle Concepts for Supersonic Applications (2001)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Investigations have been conducted in the Langley 16-Foot Transonic Tunnel (at Mach numbers from 0.60 to 1.25) and in the Langley Unitary Plan Wind Tunnel (at Mach numbers from 2.16 to 2.86) at an angle of attack of 0 deg to determine the isolated performance of several expendable nozzle concepts for supersonic nonaugmented turbojet applications. The effects of centerbody base shape, shroud length, shroud ventilation, cruciform shroud expansion ratio, and cruciform shroud flap vectoring were investigated. The nozzle pressure ratio range, which was a function of Mach number, was between 1.9 and 11.8 in the 16-Foot Transonic Tunnel and between 7.9 and 54.9 in the Unitary Plan Wind Tunnel. Discharge coefficient, thrust-minus-drag, and the forces and moments generated by vectoring the divergent shroud flaps (for Mach numbers of 0.60 to 1.25 only) of a cruciform nozzle configuration were measured. The shortest nozzle had the best thrust-minus-drag performance at Mach numbers up to 0.95 but was approached in performance by other configurations at Mach numbers of 1.15 and 1.25. At Mach numbers above 1.25, the cruciform nozzle configuration having the same expansion ratio (2.64) as the fixed geometry nozzles had the best thrust-minus-drag performance. Ventilation of the fixed geometry divergent shrouds to the nozzle external boattail flow generally improved thrust-minus-drag performance at Mach numbers from 0.60 to 1.25, but decreased performance above a Mach number of 1.25.
    Keywords: Aerodynamics
    Type: NASA/TP-2001-211259 , L-17877 , NAS 1.60:211259
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  • 3
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    High Reynolds Number Investigation of a Flush Mounted, S-Duct Inlet With Large Amounts of Boundary Layer Ingestion (2005)
    In:  Other Sources
    Details
    Publication Date: 2019-07-11
    Description: An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.This CD-ROM supplement contains inlet data including: Boundary layer data, Duct static pressure data, performance-AIP (fan face) data, Photos, Tunnel wall P-PTO data and definitions.
    Keywords: Aerodynamics
    Type: NASA/TP-2005-213766/SUPP , L-19131/SUPP
    Format: text
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  • 4
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    High Reynolds Number Investigation of a Flush-Mounted, S-Duct Inlet With Large Amounts of Boundary Layer Ingestion (2005)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.
    Keywords: Aerodynamics
    Type: NASA/TP-2005-213766 , L-19131
    Format: application/pdf
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