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  • 1
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    Ground vortex flow field investigation (1988)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Flow field investigations were conducted at the NASA Ames-Dryden Flow Visualization Facility (water tunnel) to investigate the ground effect produced by the impingement of jets from aircraft nozzles on a ground board in a STOL operation. Effects on the overall flow field with both a stationary and a moving ground board were photographed and compared with similar data found in other references. Nozzle jet impingement angles, nozzle and inlet interaction, side-by-side nozzles, nozzles in tandem, and nozzles and inlets mounted on a flat plate model were investigated. Results show that the wall jet that generates the ground effect is unsteady and the boundary between the ground vortex flow field and the free-stream flow is unsteady. Additionally, the forward projection of the ground vortex flow field with a moving ground board is one-third less than that measured over a fixed ground board. Results also showed that inlets did not alter the ground vortex flow field.
    Keywords: AERODYNAMICS
    Type: The 1987 Ground Vortex Workshop; p 61-90
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    In-flight flow visualization and pressure measurements at low speeds on the NASA F-18 high alpha research vehicle (1991)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: Inflight results from surface and off surface flow visualizations and from extensive pressure distributions document the vortical flow on the leading edge extensions (LEXs) and forebody of the NASA F-18 high alpha research vehicle (HARV) for low speeds and angles of attack up to 50 deg. Surface flow visualization data, obtained using the emitted fluid technique, were used to define separation lines and laminar separation bubbles (LSB). Off surface flow visualization data, obtained by smoke injection, were used to document both the path of the vortex cores and the location of vortex core breakdown. The location of vortex core breakdown correlated well with the loss of suction pressure on the LEX and with the flow visualization results from ground facilities. Surface flow separation lines on the LEX and forebody corresponded well with the end of pressure recovery under the vortical flows. Correlation of the pressures with wind tunnel results show fair to good correlation.
    Keywords: AERODYNAMICS
    Type: AGARD, Vortex Flow Aerodynamics; 42 p
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    In-flight flow visualization characteristics of the NASA F-18 high alpha research vehicle at high angles of attack (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Surface and off-surface flow visualization techniques were used to visualize the 3-D separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near the alpha = 25 to 26 deg and alpha = 45 to 49 deg are presented. Both the forebody and leading edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading edge flap using tufts and flow cones, and compared well with limited wind tunnel results.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4193 , H-1576 , NAS 1.15:4193
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    NASA Dryden flow visualization facility (1995)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This report describes the Flow Visualization Facility at NASA Dryden Flight Research Center, Edwards, California. This water tunnel facility is used primarily for visualizing and analyzing vortical flows on aircraft models and other shapes at high-incidence angles. The tunnel is used extensively as a low-cost, diagnostic tool to help engineers understand complex flows over aircraft and other full-scale vehicles. The facility consists primarily of a closed-circuit water tunnel with a 16- x 24-in. vertical test section. Velocity of the flow through the test section can be varied from 0 to 10 in/sec; however, 3 in/sec provides optimum velocity for the majority of flow visualization applications. This velocity corresponds to a unit Reynolds number of 23,000/ft and a turbulence level over the majority of the test section below 0.5 percent. Flow visualization techniques described here include the dye tracer, laser light sheet, and shadowgraph. Limited correlation to full-scale flight data is shown.
    Keywords: COMPUTER OPERATIONS AND HARDWARE
    Type: NASA-TM-4631 , H-1972 , NAS 1.15:4631
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Pressure distribution for the wing of the YAV-8B airplane; with and without pylons (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20 and 35 percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60 percent chord) whether or not the pylons are installed for Mach numbers greater than or equal to 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag at transonic Mach number increases.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4429 , H-1708 , NAS 1.15:4429
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Development of a low-aspect ratio fin for flight research experiments (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A second-generation flight test fixture, developed at NASA Dryden Flight Research Center, offers a generic testbed for aerodynamic and fluid mechanics research. The new fixture, a low-aspect ratio vertical fin shape mounted on the centerline of an F-15B aircraft lower fuselage, is designed for flight research at Mach numbers up to 2.0. The new fixture is a composite structure with a modular configuration and removable components for functional flexibility. This report describes the multidisciplinary design and analysis approach used to develop the fixture. The approach integrates conservative assumptions with simple analysis techniques to minimize the time and cost associated with its development. Presented are the principal disciplines required for this effort, which include aerodynamics, structures, stability, and operational considerations. In addition, preliminary results from the first phase of flight testing are presented. Acceptable directional stability and flow quality are documented and show agreement with predictions. Future envelope expansion activities will minimize current limitations so that the fixture can be used for a wide variety of high-speed aerodynamic and fluid mechanics research experiments.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4596 , H-1993 , NAS 1.15:4596 , AIAA PAPER 94-2133 , Biennial Flight Test Conference; Jun 20, 1993 - Jun 23, 1993; Colorado Springs, CO; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Description of the NASA Dryden Flow Visualization Facility (1994)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: A water tunnel facility at the NASA Dryden Flight Research Facility is described. This water tunnel facility, the Flow Visualization Facility, is used primarily for visualizing and analyzing vortical flows on aircraft models and other shapes at high incidence angles. The tunnel is used extensively as a low cost diagnostic tool to help understand complex flows over aircraft and other full scale vehicles. The facility consists primarily of a closed circuit water tunnel with a 16 in by 24 in vertical test section. The velocity of the flow through the test section can be varied from 0 to 10 in/sec, however 3 in/sec has been found to be optimum for flow visualization. This corresponds to a unit Reynolds number of 23,000 per foot and a turbulence level over the majority of the test section below .05%. Flow visualization techniques described include the dye tracer, laser light sheet and the shadowgraph techniques. Limited correlation to full-scale flight data is shown.
    Keywords: Research and Support Facilities (Air)
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    Global Test Range: Toward Airborne Sensor Webs (2008)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This viewgraph presentation reviews the planned global sensor network that will monitor the Earth's climate, and resources using airborne sensor systems. The vision is an intelligent, affordable Earth Observation System. Global Test Range is a lab developing trustworthy services for airborne instruments - a specialized Internet Service Provider. There is discussion of several current and planned missions.
    Keywords: Earth Resources and Remote Sensing
    Type: W-HALES 2008: NASA-NICT Joint Workshop on HALE UAV and Wireless Systems; Apr 17, 2008 - Apr 18, 2008; Honolulu, HI; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    In-flight flow visualization results from the X-29A aircraft at high angles of attack (1992)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Flow visualization techniques were used on the X-29A aircraft at high angles of attack to study the vortical flow off the forebody and the surface flow on the wing and tail. The forebody vortex system was studied because asymmetries in the vortex system were suspected of inducing uncommanded yawing moments at zero sideslip. Smoke enabled visualization of the vortex system and correlation of its orientation with flight yawing moment data. Good agreement was found between vortex system asymmetries and the occurrence of yawing moments. Surface flow on the forward-swept wing of the X-29A was studied using tufts and flow cones. As angle of attack increased, separated flow initiated at the root and spread outboard encompassing the full wing by 30 deg angle of attack. In general, the progression of the separated flow correlated well with subscale model lift data. Surface flow on the vertical tail was also studied using tufts and flow cones. As angle of attack increased, separated flow initiated at the root and spread upward. The area of separated flow on the vertical tail at angles of attack greater than 20 deg correlated well with the marked decrease in aircraft directional stability.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4430 , H-1825 , NAS 1.15:4430 , AIAA PAPER 92-4102 , Biennial Flight Test Conference; Aug 24, 1992 - Aug 26, 1992; Hilton Head, SC; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    A summary of the forebody high-angle-of-attack aerodynamics research on the F-18 and the X-29A aircraft (1992)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: High-angle-of-attack aerodynamic studies have been conducted on both the F18 High Alpha Research Vehicle (HARV) and the X-29A aircraft. Data obtained include on- and off-surface flow visualization and static pressure measurements on the forebody. Comparisons of similar results are made between the two aircraft where possible. The forebody shapes of the two aircraft are different and the X-29A forebody flow is affected by the addition of nose strakes and a flight test noseboom. The forebody flow field of the F-18 HARV is fairly symmetric at zero sideslip and has distinct, well-defined vortices. The X-29A forebody vortices are more diffuse and are sometimes asymmetric at zero sideslip. These asymmetries correlate with observed zero-sideslip aircraft yawing moments.
    Keywords: AERODYNAMICS
    Type: NASA-TM-104261 , H-1862 , NAS 1.15:104261 , SAE-92-1996 , SAE Aerotech 1992 Conference; Oct 05, 1992 - Oct 08, 1992; Anaheim, CA; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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