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  • 1
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    Effect of Boattail and Sidewall Curvature on Nozzle Drag Characteristics (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The NASA-industry team has sponsored several studies in the last two years to address the installed nozzle boattail drag issues. Some early studies suggested that nozzle boattail drag could be as much as 25 to 40 percent of the subsonic cruise. As part of this study tests have been conducted at NASA-Langley to determine the uninstalled drag characteristics of a proposed nozzle. The overall objective was to determine the effects of nozzle external flap curvature and sidewall boattail variations. This test would also provide data for validating CFD predictions of nozzle boattail drag.
    Keywords: Aerodynamics
    Type: 1997 NASA High-Speed Research Program Aerodynamic Performance Workshop; Volume 1; Part 1; 669-706; NASA/CP-1999-209691/VOL1/PT1
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Effect of Boattail and Sidewall Curvature on Nozzle Drag Characteristics (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The NASA-industry team has sponsored several studies in the last two years to address the installed nozzle boattail drag issues. Some early studies suggested that nozzle boattail drag could be as much as 25 to 40 percent of the subsonic cruise. As part of this study tests have been conducted at NASA-Langley to determine the uninstalled drag characteristics of a proposed nozzle. The overall objective was to determine the effects of nozzle external flap curvature and sidewall boattail variations. This test would also provide data for validating CFD predictions of nozzle boattail drag.
    Keywords: Aerodynamics
    Type: 1997 NASA High-Speed Research Program Aerodynamic Performance Workshop; Volume 1; Part 1; 669-706; NASA/CP-1999-209691/VOL1/PT1
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Aftbody Closure Model Design: Lessons Learned (1999)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: An Aftbody Closure Test Program is necessary in order to provide aftbody drag increments that can be added to the drag polars produced by testing the performance models (models 2a and 2b). These models had a truncated fuselage, thus, drag was measured for an incomplete configuration. In addition, trim characteristics cannot be determined with a model with a truncated fuselage. The stability and control tests were conducted with a model (model 20) having a flared aftbody. This type aftbody was needed in order to provide additional clearance between the base of the model and the sting. This was necessary because the high loads imposed on the model for stability and control tests result in large model deflections. For this case, the aftbody model will be used to validate stability and control performance.
    Keywords: Aircraft Design, Testing and Performance
    Type: 1998 NASA High-Speed Research Program Aerodynamic Performance Workshop; Volume 1; Part 2; 1545-1568; NASA/CP-1999-209692/VOL1/PT2
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Multiaxis Thrust-Vectoring Characteristics of a Model Representative of the F-18 High-Alpha Research Vehicle at Angles of Attack From 0 deg to 70 deg (1995)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the multiaxis thrust-vectoring characteristics of the F-18 High-Alpha Research Vehicle (HARV). A wingtip supported, partially metric, 0.10-scale jet-effects model of an F-18 prototype aircraft was modified with hardware to simulate the thrust-vectoring control system of the HARV. Testing was conducted at free-stream Mach numbers ranging from 0.30 to 0.70, at angles of attack from O' to 70', and at nozzle pressure ratios from 1.0 to approximately 5.0. Results indicate that the thrust-vectoring control system of the HARV can successfully generate multiaxis thrust-vectoring forces and moments. During vectoring, resultant thrust vector angles were always less than the corresponding geometric vane deflection angle and were accompanied by large thrust losses. Significant external flow effects that were dependent on Mach number and angle of attack were noted during vectoring operation. Comparisons of the aerodynamic and propulsive control capabilities of the HARV configuration indicate that substantial gains in controllability are provided by the multiaxis thrust-vectoring control system.
    Keywords: Aerodynamics
    Type: NASA-TP-3531 , L-17441 , NAS 1.60:3531
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    The NASA Langley 16-Foot Transonic Tunnel: Historical Overview, Facility Description, Calibration, Flow Characteristics, and Test Capabilities (1995)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Langley 16-Foot Transonic Tunnel is a closed-circuit single-return atmospheric wind tunnel that has a slotted octagonal test section with continuous air exchange. The wind tunnel speed can be varied continuously over a Mach number range from 0.1 to 1.3. Test-section plenum suction is used for speeds above a Mach number of 1.05. Over a period of some 40 years, the wind tunnel has undergone many modifications. During the modifications completed in 1990, a new model support system that increased blockage, new fan blades, a catcher screen for the first set of turning vanes, and process controllers for tunnel speed, model attitude, and jet flow for powered models were installed. This report presents a complete description of the Langley 16-Foot Transonic Tunnel and auxiliary equipment, the calibration procedures, and the results of the 1977 and the 1990 wind tunnel calibration with test section air removal. Comparisons with previous calibrations showed that the modifications made to the wind tunnel had little or no effect on the aerodynamic characteristics of the tunnel. Information required for planning experimental investigations and the use of test hardware and model support systems is also provided.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: NASA-TP-3521 , NAS 1.60:3521 , L-17445
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Static internal performance of an axisymmetric nozzle with multiaxis thrust-vectoring capability (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the static test facility of the Langley 16 Foot Transonic Tunnel in order to determine the internal performance characteristics of a multiaxis thrust vectoring axisymmetric nozzle. Thrust vectoring for this nozzle was achieved by deflection of only the divergent section of this nozzle. The effects of nozzle power setting and divergent flap length were studied at nozzle deflection angles of 0 to 30 at nozzle pressure ratios up to 8.0.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4237 , L-16809 , NAS 1.15:4237
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    A static investigation of the thrust vectoring system of the F/A-18 high-alpha research vehicle (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A static (wind-off) test was conducted in the static test facility of the Langley 16-foot Transonic Tunnel to evaluate the vectoring capability and isolated nozzle performance of the proposed thrust vectoring system of the F/A-18 high alpha research vehicle (HARV). The thrust vectoring system consisted of three asymmetrically spaced vanes installed externally on a single test nozzle. Two nozzle configurations were tested: A maximum afterburner-power nozzle and a military-power nozzle. Vane size and vane actuation geometry were investigated, and an extensive matrix of vane deflection angles was tested. The nozzle pressure ratios ranged from two to six. The results indicate that the three vane system can successfully generate multiaxis (pitch and yaw) thrust vectoring. However, large resultant vector angles incurred large thrust losses. Resultant vector angles were always lower than the vane deflection angles. The maximum thrust vectoring angles achieved for the military-power nozzle were larger than the angles achieved for the maximum afterburner-power nozzle.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4359 , L-17002 , NAS 1.15:4359
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Study of external dynamic flap loads on a 6 percent B-1B model (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The origin of dynamic pressure loads on external divergent engine nozzle flaps of the B-1B aircraft was investigated in the NASA/LaRC 16 foot transonic tunnel using a 6 percent full span model with powered engine nacelles. External flap dynamic loads and afterbody drag associated with flap removal were measured using this model. Both dry and max. A/B power nozzles were evaluated in this study. As a result of this study, the principal mechanisms responsible for high dynamic external flap loads were determined along with performance penalty associated with flap removal.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: ASME PAPER 91-GT-236
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    Parametric investigation of single-expansion-ramp nozzles at Mach numbers from 0.60 to 1.20 (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of varying six nozzle geometric parameters on the internal and aeropropulsive performance characteristics of single-expansion-ramp nozzles. This investigation was conducted at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.5 to 12, and angles of attack of 0 deg +/- 6 deg. Maximum aeropropulsive performance at a particular Mach number was highly dependent on the operating nozzle pressure ratio. For example, as the nozzle upper ramp length or angle increased, some nozzles had higher performance at a Mach number of 0.90 because of the nozzle design pressure was the same as the operating pressure ratio. Thus, selection of the various nozzle geometric parameters should be based on the mission requirements of the aircraft. A combination of large upper ramp and large lower flap boattail angles produced greater nozzle drag coefficients at Mach number greater than 0.80, primarily from shock-induced separation on the lower flap of the nozzle. A static conditions, the convergent nozzle had high and nearly constant values of resultant thrust ratio over the entire range of nozzle pressure ratios tested. However, these nozzles had much lower aeropropulsive performance than the convergent-divergent nozzle at Mach number greater than 0.60.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3240 , L-17067 , NAS 1.60:3240
    Format: application/pdf
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  • 10
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    Computations for the 16-foot transonic tunnel, NASA, Langley Research Center, revision 1 (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The equations used by the 16 foot transonic tunnel in the data reduction programs are presented in eight modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: tunnel parameters; jet exhaust measurements; skin friction drag; balance loads and model attitudes calculations; internal drag (or exit-flow distributions); pressure coefficients and integrated forces; thrust removal options; and turboprop options. This document is a companion document to NASA TM-83186, A User's Guide to the Langley 16 Foot Transonic Tunnel, August 1981.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: NASA-TM-86319-REV-1 , NAS 1.15:86319-REV-1
    Format: application/pdf
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