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  • 1
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    Effect of underwing aft-mounted nacelles on the longitudinal aerodynamic characteristics of a high-wing transport airplane (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: As part of a propulsion/airframe integration program, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the longitudinal aerodynamic effects of installing flow through engine nacelles in the aft underwing position of a high wing transonic transfer airplane. Mixed flow nacelles with circular and D-shaped inlets were tested at free stream Mach numbers from 0.70 to 0.85 and angles of attack from -2.5 deg to 4.0 deg. The aerodynamic effects of installing antishock bodies on the wing and nacelle upper surfaces as a means of attaching and supporting nacelles in an extreme aft position were investigated.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2447 , L-15664 , NAS 1.60:2447
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Integration effects of underwing forward- and rearward-mounted separate-flow, flow-through nacelles on a high-wing transport (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted in the Langley 16-Foot Transonic Tunnel at free-stream Mach numbers from 0.70 to 0.82 and angles of attack from -2.5 to 4.0 degrees to determine the integration effects of pylon-mounted underwing forward and rearward separate-flow, flow-through nacelles on a high-wing transonic transport configuration. The results showed that the installed drag of the nacelle/pylon in the rearward location was slightly less than that of the nacelle/pylon in the forward location. This reduction was due to the reduction in calculated skin friction of the nacelle/pylon configuration. In all cases the combined value of form, wave, and interference drag was excessively high. However, the configuration with the nacelle/pylon in a rearward location produced an increase in lift over that of the basic wing-body configuration.
    Keywords: AERODYNAMICS
    Type: NASA-TM-87627 , L-16026 , NAS 1.15:87627
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Aerodynamic characteristics of a high-wing transport configuration with a over-the-wing nacelle-pylon arrangement (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects on the aerodynamic characteristics of a high-wing transport configuration of installing an over-the-wing nacelle-pylon arrangement. The tests are conducted at Mach numbers from 0.70 to 0.82 and at angles of attack from -2 deg to 4 deg. The configurational variables under study include symmetrical and contoured nacelles and pylons, pylon size, and wing leading-edge extensions. The symmetrical nacelles and pylons reduce the lift coefficient, increase the drag coefficient, and cause a nose-up pitching-moment coefficient. The contoured nacelles significantly reduce the interference drag, though it is still excessive. Increasing the pylon size reduces the drag, whereas adding wing leading-edge extension does not affect the aerodynamic characteristics significantly.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2497 , L-15959 , NAS 1.60:2497
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    16-foot transonic tunnel test section flowfield survey (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A flow survey has been made of the test section of the NASA Langley Research Center 16-Foot Transonic Tunnel at subsonic and supersonic speeds. The survey was performed using five five-hole pyramid-head probes mounted at 14 inch intervals on a survey rake. Probes were calibrated at freestream Mach numbers from 0.50 to 0.95 and from 1.18 to 1.23. Flowfield surveys were made at Mach numbers from 0.50 to 0.90 and at Mach 1.20. The surveys were made at tunnel stations 130.6, 133.6, and 136.0. By rotating the survey rake through 180 degrees, a cylindrical volume of the test section 4.7 feet in diameter and 5.4 feet long centered about the tunnel centerline was surveyed. Survey results showing the measured test section upflow and sideflow characteristics and local Mach number distributions are presented. The report documents the survey probe calibration techniques used, summarizes the procedural problems encountered during testing, and identifies the data discrepancies observed during the post-test data analysis.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: NASA-TM-109157 , NAS 1.15:109157
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Investigation of the flow field surrounding circular-arc boattail nozzles at subsonic speeds (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of jet exhaust on the subsonic flow field surrounding boattail nozzles with attached and separated boundary layers were investigated. Measurements of local Mach numbers and flow angles were made at free-stream Mach numbers of 0.60 and 0.80 at an angle of attack of 0 deg. Jet exhaust flow was simulated with a solid cylindrical sting and with high pressure air at jet-nozzle total pressure ratios of 2.9 and 5.0. Results show strong effects of the jet-wave structure on the external flow field. The predicted local Mach numbers and flow angles for attached-flow nozzles with solid jet simulators obtained by using subsonic inviscid/viscous-flow theory are in good agreement with experimental data. Prediction of nozzle surface pressure distributions which include jet-entrainment effects also agree with experimental data for attached-flow nozzles with high pressure air jets.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1633 , L-13318
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Boundary-layer separation on isolated boattail nozzles (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An angle of attack of 0 deg was investigated in the Langley 16 foot transonic tunnel at free-stream Mach numbers from 0.40 to 0.95 to study the phenomenon of separated flow on a series of circular-arc afterbodies. Both high-pressure air and solid circular cylinders with the cylinder diameter equal to the nozzle-exit diameter were used to simulate jet exhausts. The results indicate that boundary-layer separation is most extensive on steep boattails at high Mach numbers. The jet total-pressure ratio changes (jet total pressure to free-stream static pressure) affected the extent of separation very little; however, comparison of the separation data obtained by using the two jet-simulation techniques indicate that entrainment associated with the presence of a jet had a significant effect on the extent of separation. The solid-simulator separation data were also used to evaluate the predictions of eight separation criteria.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1226 , L-12104
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Boundary layer separation on isolated boattail nozzles (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The phenomenon of separated flow on a series of circular-arc afterbodies was investigated using the Langley 16-foot transonic tunnel at free-stream Mach numbers from 0.40 to 0.95 at 0 deg angle of attack. Both high-pressure air and solid circular cylinders with a diameter equal to the nozzle exit diameter were used to simulate jet exhausts. A detailed data base of boundary layer separation locations was obtained using oil-flow techniques. The results indicate that boundary layer separation is most extensive on steep boattails at high Mach numbers.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152703
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Experimental and theoretical study of flow fields surrounding boattail nozzles at subsonic speeds (1976)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: An investigation has been conducted to determine the effects of jet exhaust on the subsonic flow surrounding boattail nozzles with and without separated boundary layers. Measurements of local Mach number and flow angle were made at free-stream Mach numbers of 0.60 and 0.80. Jet exhaust flow was simulated with a solid sting and with high-pressure air at jet total pressure ratios of 3 and 5. The results show that there are strong effects of the wave structure of the jet on the external flow field. The local Mach numbers and flow angles for the models with solid sting jet simulators were in good agreement with subsonic inviscid/viscous flow theory. The theoretical method does not, however, adequately predict the flow about the models with high-pressure air jets.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 76-675 , Propulsion Conference; Jul 26, 1976 - Jul 29, 1976; Palo Alto, CA
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    Experimental investigation of two nonaxisymmetric wedge nozzles at free stream Mach numbers up to 1.20 (1982)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Forces and pressures on two nonaxisymmetric wedge nozzles were measured in a 16 foot transonic tunnel. Tests were conducted at static conditions and at free stream Mach numbers of 0.60, 0.80, 0.90, 0.94, and 1.20. The range of nozzle pressure ratios varied with configuration and Mach number. The internal and external geometry of the nozzles and the test model are defined in detail. Nozzle performance data are presented as discharge coefficients, internal thrust ratios, thrust minus nozzle drag ratios, and ideal thrust coefficients. Extensive internal and external pressure measurements are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2054 , L-15276 , NAS 1.60:2054
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    NASA TECHNICAL REPORTS
  • 10
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    Transonic perturbation analysis of wing-fuselage-nacelle-pylon configurations with powered jet exhausts (1982)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A transonic small perturbation method has been developed for the analysis of general wing-fuselage-nacelle-pylon configurations with powered jet exhausts. Finite difference successive line relaxation algorithm is used to solve the small disturbance potential equation in conservative form. The nacelle tangency condition and the jet exhaust plume contact conditions are fulfilled in a quasi-cylindrical fashion on a surface fitting the Cartesian grid. The pylon tangency condition is treated in a quasi-planar manner as for the wing. Viscous displacement effects on the wing are modeled by suitable shape changes including the placement of a viscous ramp at the base of the shock. Computed results of a transport configuration show satisfactory correlation with test data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 82-0255 , Aerospace Sciences Meeting; Jan 11, 1982 - Jan 14, 1982; Orlando, FL
    Format: text
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    NASA TECHNICAL REPORTS
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