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  • 1
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    Effect of Sweep on Cavity Flow Fields at Subsonic and Transonic Speeds (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An experimental investigation was conducted in the NASA Langley 7 x 10-Foot High Speed Tunnel (HST) to study the effect of leading- and trailing-edge sweep on cavity flow fields for a range of cavity length-to-height (l/h) ratios. The free-stream Mach number was varied from 0.2 to 0.8. The cavity had a depth of 0.5 inches, a width of 2.5 inches, and a maximum length of 12.0 inches. The leading- and trailing-edge sweep was adjusted using block inserts to achieve leading edge sweep angles of 65 deg, 55 deg, 45 deg, 35 deg, and 0 deg. The fore and aft cavity walls were always parallel. The aft wall of the cavity was remotely positioned to achieve a range of length-to-depth ratios. Fluctuating- and static-pressure data were obtained on the floor of the cavity. The fluctuating pressure data were used to determine whether or not resonance occurred in the cavity rather than to provide a characterization of the fluctuating pressure field. Qualitative surface flow visualization was obtained using a technique in which colored water was introduced into the model through static-pressure orifices. A complete tabulation of the mean static-pressure data for the swept leading edge cavities is included.
    Keywords: Aerodynamics
    Type: NASA/TM-2012-217577 , NF1676L-13087 , L-20047
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Experimental Measurements of Sonic Boom Signatures Using a Continuous Data Acquisition Technique (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A wind tunnel investigation was conducted in the Langley Unitary Plan Wind Tunnel to determine the effectiveness of a technique to measure aircraft sonic boom signatures using a single conical survey probe while continuously moving the model past the probe. Sonic boom signatures were obtained using both move-pause and continuous data acquisition methods for comparison. The test was conducted using a generic business jet model at a constant angle of attack and a single model-to-survey-probe separation distance. The sonic boom signatures were obtained at a Mach number of 2.0 and a unit Reynolds number of 2 million per foot. The test results showed that it is possible to obtain sonic boom signatures while continuously moving the model and that the time required to acquire the signature is at least 10 times faster than the move-pause method. Data plots are presented with a discussion of the results. No tabulated data or flow visualization photographs are included.
    Keywords: Aerodynamics
    Type: NASA/TP-2013-218035 , L-20226 , NF1676L-16049
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Effectiveness of a Wedge Probe to Measure Sonic Boom Signatures in a Supersonic Wind Tunnel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A wind tunnel investigation was conducted in the Langley Unitary Plan Wind Tunnel (UPWT) to determine the effectiveness of a wedge probe to measure sonic boom pressure signatures compared to a slender conical probe. A generic business jet model at a constant angle of attack and at a single model to probe separation distance was used to generate a sonic boom signature. Pressure signature data were acquired with both the wedge probe and a slender conical probe for comparison. The test was conducted at a Mach number of 2.0 and a free-stream unit Reynolds number of 2 million per foot. The results showed that the wedge probe was not effective in measuring the sonic boom pressure signature of the aircraft model in the supersonic wind tunnel. Data plots and a discussion of the results are presented. No tabulated data or flow visualization photographs are included.
    Keywords: Aerodynamics
    Type: NASA/TP-2013-218036 , L-20239 , NF1676L-16253
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    NASA TECHNICAL REPORTS
  • 4
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    Experimental Sonic Boom Measurements on a Mach 1.6 Cruise Low-Boom Configuration (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A wind tunnel test has been conducted by Gulfstream Aerospace Corporation (GAC) to measure the sonic boom pressure signature of a low boom Mach 1.6 cruise business jet in the Langley Unitary Plan Wind Tunnel at Mach numbers 1.60 and 1.80. Through a cooperative agreement between GAC and the National Aeronautics and Space Administration (NASA), GAC provided NASA access to some of the experimental data and NASA is publishing these data for the sonic boom research community. On-track and off-track near field sonic boom pressure signatures were acquired at three separation distances (0.5, 1.2, and 1.7 reference body lengths) and three angles of attack (-0.26deg, 0.26deg, and 0.68deg). The model was blade mounted to minimize the sting effects on the sonic boom signatures. Although no extensive data analysis is provided, selected data are plotted to illustrate salient features of the data. All of the experimental sonic boom pressure data are tabulated. Schlieren images of the configuration are also included.
    Keywords: Aerodynamics
    Type: NASA/TM-2012-217598 , L-20103 , NF1676L-13939
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Vortex Flows at Supersonic Speeds (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A review of research conducted at the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data are for flat plates, cavities, bodies, missiles, wings, and aircraft with Mach numbers of 1.5 to 4.6. Data are presented to show the types of vortex structures that occur at supersonic speeds and the impact of these flow structures on vehicle performance and control. The data show the presence of both small- and large-scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices. Data are shown that highlight the effect of leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber on the aerodynamics of and flow over delta wings. Finally, a discussion of a design approach for wings that use vortex flows for improved aerodynamic performance at supersonic speeds is presented.
    Keywords: Aerodynamics
    Type: NASA/TP-2003-211950 , L-18008 , NAS 1.60:211950
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    ARES I Aerodynamic Testing at the NASA Langley Unitary Plan Wind Tunnel (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Small-scale force and moment and pressure models based on the outer mold lines of the Ares I design analysis cycle crew launch vehicle were tested in the NASA Langley Research Center Unitary Plan Wind Tunnel from May 2006 to September 2009. The test objectives were to establish supersonic ascent aerodynamic databases and to obtain force and moment, surface pressure, and longitudinal line-load distributions for comparison to computational predictions. Test data were obtained at low through high supersonic Mach numbers for ranges of the Reynolds number, angle of attack, and roll angle. This paper focuses on (1) the sensitivity of the supersonic aerodynamic characteristics to selected protuberances, outer mold line changes, and wind tunnel boundary layer transition techniques, (2) comparisons of experimental data to computational predictions, and (3) data reproducibility. The experimental data obtained in the Unitary Plan Wind Tunnel captured the effects of evolutionary changes to the Ares I crew launch vehicle, exhibited good agreement with predictions, and displayed satisfactory within-test and tunnel-to-tunnel data reproducibility.
    Keywords: Aerodynamics
    Type: AIAA Paper 2011-0999 , NF1676L-10544 , 49th AIAA Aerospace Sciences Meeting; Jan 04, 2011 - Jan 07, 2011; Orlando, FL; United States
    Format: application/pdf
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