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  • 1
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    Comparison of algebraic turbulence models for afterbody flows with jet exhaust (1992)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Description: Three-dimensional Navier-Stokes simulations have been done for transonic and low supersonic flow past a nonaxisymmetric nozzle typical of those advocated for advanced fighter airplanes. The jet exhaust is included in the calculations. The investigation compares the performance of the unmodified Baldwin-Lomax turbulence model with its performance when enhanced by the Degani-Schiff and the Goldberg modifications. Solutions are presented for Mach numbers of 0.80, 0.94, and 1.20 at 0-deg angle of attack and a Reynolds number of 20 x 10 to the 6th. The numerical results, which are compared to the wind-tunnel data, show that the three turbulence models predict considerably different shock locations, separated-flow regions, and flowfields.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 30; 11; p. 2716-2722.
    Format: text
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  • 2
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    Natural laminar flow nacelle for transport aircraft (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The potential of laminar flow nacelles for reducing installed engine/nacelle drag was studied. The purpose was twofold: to experimentally verify a method for designing laminar flow nacelles and to determine the effect of installation on the extent of laminar flow on the nacelle and on the nacelle pressure distributions. The results of the isolated nacelle tests illustrated that laminar flow could be maintained over the desired length. Installing the nacelles on wing/pylon did not alter the extent of laminar flow occurring on the nacelles. The results illustrated that a significant drag reduction was achieved with this laminar flow design. Further drag reduction could be obtained with proper nacelle location and pylon contouring.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Langley Symposium on Aerodynamics, Volume 1; p 445-460
    Format: application/pdf
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  • 3
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    A Wind Tunnel Investigation of Three NACA 1-Series Inlets at Mach Numbers up to 0.92 (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Pressure distributions on three NACA 1-series inlets have been obtained in the Langley 16-Foot Transonic Tunnel. The cowl diameter ratio (ratio of cowl highlight diameter to cowl maximum diameter) was 0.85 for all three inlets. The cowl length ratio (ratio of cowl length to cowl maximum diameter) was 1.0 for two of the inlets (NACA 1-85-100) and 0.439 for the other (NACA 1-85-43.9) inlet. One of the inlets with a cowl length ratio of 1.0 had an internal contraction ratio (ratio of highlight area to throat area) of 1.009 and the other had a contraction ratio of 1.250. The inlet with a cowl length ratio of 0.439 also had an internal contraction ratio of 1.250. All three inlets had longitudinal rows of static pressure orifices on the top and bottom external cowl surfaces. The inlet with a contraction ratio of 1.009 also had a row of static pressure orifices on the side of the cowl (external surface). The two inlets with a contraction ratio of 1.250 had a longitudinal row of static pressure orifices on the diffuser surface.
    Keywords: Aerodynamics
    Type: NASA-TM-110300 , NAS 1.15:110300
    Format: application/pdf
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  • 4
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    Effect of location of aft-mounted nacelles on the longitudinal aerodynamic characteristics of a high-wing transport airplane (1990)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: As part of a propulsion/airframe integration program at Langley Research Center, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of locating flow-through mixed flow engine nacelles in several aft underwing positions on the longitudinal aerodynamics of a high wing transport airplane. D-shaped inlet nacelles were used in the test. Some configurations with antishock bodies and with nacelle toe-in were also tested. Data were obtained for a free stream Mach number range of 0.70 to 0.85 and a model angle-of-attack range from -2.5 to 4.0 degrees.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3047 , L-16743 , NAS 1.60:3047
    Format: application/pdf
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  • 5
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    Wind tunnel investigation of three axisymmetric cowls of different lengths at Mach numbers from 0.60 to 0.92 (1993)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Pressure distributions on three inlets having different cowl lengths were obtained in the Langley 16-Foot Transonic Tunnel. The cowl diameter ratio (highlight diameter to maximum diameter) was 0.85 and the cowl length ratios (cowl length to maximum diameter) were 0.337, 0.439, and 0.547. The cowls had identical nondimensionalized (with respect to cowl length) external geometry and identical internal geometry. The internal contraction ratio (highlight area to throat area) was 1.250. The inlets had longitudinal rows of static pressure orifices on the top and bottom (external) surfaces and on the contraction (internal) and diffuser surfaces. The afterbody was cylindrical in shape, and its diameter was equal to the maximum diameter of the cowl. Depending on the cowl configuration and free-stream Mach number, the mass-flow ratio varied between 0.27 and 0.87 during the tests. Angle of attack varied from 0 to 4.1 deg at selected Mach numbers and mass-flow ratios, and the Reynolds number varied with the Mach number from 3.2x10(exp 6) to 4.2x10(exp 6) per foot.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4488 , L-17191 , NAS 1.15:4488
    Format: application/pdf
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  • 6
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    Wingtip vortex turbine investigation for vortex energy recovery (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A flight test investigation has been conducted to determine the performance of wingtip vortex turbines and their effect on aircraft performance. The turbines were designed to recover part of the large energy loss (induced drag) caused by the wingtip vortex. The turbine, driven by the vortex flow, reduces the strength of the vortex, resulting in an associated induced drag reduction. A four-blade turbine was mounted on each wingtip of a single-engine, T-tail, general aviation airplane. Two sets of turbine blades were tested, one with a 15' twist (washin) and one with no twist. Th power recovered by the turbine and the installed drag increment were measured. A trade-off between turbine power and induced drag reduction was found to be a function of turbine blade incidence angle. This test has demonstrated that the wingtip vortex turbine is an attractive alternate, as well as an emergency, power source.
    Keywords: AERODYNAMICS
    Type: SAE PAPER 901936
    Format: text
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  • 7
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    Transonic Navier-Stokes solutions of three-dimensional afterbody flows (1989)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The performance of a three-dimensional Navier-Stokes solution technique in predicting the transonic flow past a nonaxisymmetric nozzle was investigated. The investigation was conducted at free-stream Mach numbers ranging from 0.60 to 0.94 and an angle of attack of 0 degrees. The numerical solution procedure employs the three-dimensional, unsteady, Reynolds-averaged Navier-Stokes equations written in strong conservation form, a thin layer assumption, and the Baldwin-Lomax turbulence model. The equations are solved by using the finite-volume principle in conjunction with an approximately factored upwind-biased numerical algorithm. In the numerical procedure, the jet exhaust is represented by a solid sting. Wind-tunnel data with the jet exhaust simulated by high pressure air were also obtained to compare with the numerical calculations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4111 , L-16516 , NAS 1.15:4111
    Format: application/pdf
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  • 8
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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
    Format: application/pdf
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