Langley Research Center Turbulence Modeling Resource

Exp: NASA Juncture Flow (JF) - Transitional Symmetric Wing Geometry

Return to: Exp: NASA Juncture Flow - Intro Page for Transitional Symmetric Wing

CAD FILES FOR THE SYMMETRIC CONFIGURATION (WITH HORN)

The CAD files are the same for the turbulent symmetric wing and transitional symmetric wing.

• As-designed symmetric wing and body with horn/leading edge extension:
  • Parasolid format (units of meters)
  • STEP format (units of mm)
  • IGES format (units of mm)
  
   
• In_tunnel_symmetric.tar.gz (gzipped tar file, 10.3 MB)
  • Contains 14x22 tunnel definition, as-designed symmetric wing and body with horn/leading edge extension placed with respect to the tunnel at 3 different angles of incidence (0, 1, and 5 deg), and approximated mast/sting definitions).
  • Some important information needed for running Juncture Flow CFD in the tunnel is provided on the JF Turbulent Symmetric Wing Boundary Conditions page.

Notes:

1. Three CAD formats are provided: The Parasolid (x_t) files are the "native" files, with default units of meters. The STEP (stp) and IGES (igs) file versions are given in mm. Because the experimental data are provided in mm, for direct comparisons it is recommended that units of mm be used when building CFD grids.
2. The experimental data has the fuselage nose at (0,0,0), x axis along the body axis, z up, and y out the starboard side. Also, all LDV data profiles (as well as all surface pressure taps) are aligned with this body axis system. When extracting velocities, Reynolds stresses, etc. from CFD, this is the coordinate system that you need to use when comparing with the experimental data. The "as-designed symmetric wing and body" CAD definitions are already in this coordinate system.
3. None of the files provided above account for any aeroelastic deflection.
4. The CAD of the wind tunnel high-speed leg was created from a laser scan, but was simplified (many recesses etc were faired over or ignored). This simplified as-built model was also used/evaluated in AIAA Paper 2015-2022. The wind tunnel CAD is positioned with x=0 at the start of the test section, and y=z=0 is along the "centerline" (z=up).
5. The wind tunnel's outer wall (sometimes referred to as the "north wall") is located to the left when facing forward (pilot's view); the inner wall ("south wall") is located to the right.
6. The wind tunnel has a 127 mm wide flow-straightening honeycomb centered near x = -20.117 m, followed by 4 screens near x = -19.291, -19.063, -18.885, and -18.707 m. For this and other 14x22 wind tunnel details, see NASA TP 3008, September 1990 and NASA TM 85662, December 1983. (Note the latter paper title refers to the 14x22 tunnel in metric units: 4- by 7-Meter. However, nowadays the tunnel is typically referred to by its name referenced in feet.)
7. In the wind tunnel, the approximate locations of the probes used to determine running conditions are:
   • Total Pressure: (x,y,z) = (-18.105, -7.010, -4.084) m (1.646 m away from the north wall)
   • Static Pressure: (x,y,z) = (-3.658, -3.048, 0) m (355.60 mm away from the north wall)
   • Stagnation Temperature: (x,y,z) = (-0.140, -3.234, -0.269) m (82.55 mm away from the north wall)
   • Dewpoint sensor orifice: (x,y,z) = (-0.140, -3.317, -0.187) m (on the north wall)
8. The symmetric wing is a unique blend of NACA 0015 shape near the root, with other (thinner) NACA 0012 and 0010 shapes outboard. The horn (leading edge fillet) at the wing root leading edge of course alters the wing shape inboard. The fuselage used in the Juncture Flow experiment is the same for all tests.
9. Data were not taken to determine angle of attack corrections (accounting for influence of the tunnel walls). Therefore, in free-air CFD runs to date, nominal incidence angles were used for freestream angle of attack. This is obviously not strictly correct, but arguably the effect on many quantities of interest (QoI) is relatively minor.
10. In the experiment, forces and moments were not measured.
11. During the tunnel runs, the goal was to try to keep a reference point on the model (2.448 m behind the model nose tip and on the fuselage centerline) located approximately 5.4 m from the test section entrance and at the center of the test section (approx 2.2098 m above the floor). However, at some negative angles, this was impossible. The equations for the height of the model's reference point (x=2448 mm, z=0 mm) above the floor can be approximately expressed as:
    • For -10 ≤ alpha < -2.75, H = 2540.1 + 117.91*alpha (mm)
    • For -2.75 ≤ alpha ≤ 10, H = 2209.8 mm
12. The roll angle was never intentionally adjusted but did show a roughly linear relationship with the angle of incidence, ranging from approximately -0.07 deg at pitch angle = 10 deg to 0.07 deg at -10 deg.
13. The average roughness and the rms roughness for the painted model surface was 2.70 +- 0.94 μm and 3.33 +- 1.12 μm, respectively.

DISCUSSION OF MODEL TRIPS

Some of the transition test was performed with boundary layer tripping on the fuselage nose, and some without. (There were no trips on the wings for the transition test.) For fuselage tripping, 11.4 mil (289.4 micrometer) high trip dots (blue color) were placed on the fuselage at an arc distance of approximately 16 inches (406 mm) from the fuselage nose, or at an x location of approximately x=336 mm. The trips were cylindrical with diameter 1.16 mm and center-to-center spacing of 2.47 mm. Their effectiveness was verified by infrared thermography.

For the trip dots placed by hand around the nose of the fuselage, nominally the goal was to place them on the side of the fuselage a specific distance from the nose (running along the surface), then to maintain this x-location all the way around. From the laser scan, the actual x-location varied in the range 324 < x < 359 mm, approximately as follows:

• On the flat port side: x=329 mm top flat section, increasing to x=339 mm bottom flat section
• On the flat starboard side: x=340 mm top flat section, increasing to x=343 mm bottom flat section
• On the top: x=329 mm left, decreasing to 324 mm near y=-112 mm, then increasing to x=340 mm right
• On the bottom: x=339 mm left, increasing to x=359 mm near y=-8 mm, then decreasing to x=343 mm right

The following figure is from an STL file created from the scanned points of the symmetric wing and body with horn/leading edge extension (note that the STL file introduces additional wiggles and non-smoothness that is not present in the real geometry). This scan result can be used to visualize and verify the trip dot placement.

 

DISCUSSION OF MODEL DEFORMATION

There are always aeroelastic deflections in any model. These are not accounted for in any of the files provided above. Details for the symmetric wing are not currently available. However, see the Turbulent F6-Based Geometry page for discussion of photogrammetry results for the F6-based wing.
 

DISCUSSION OF AS-BUILT VS. AS-DESIGNED SHAPE OF THE MODEL

A detailed discussion of the as-built shape of the symmetric model is provided on the page: Turbulent Symmetric Wing Geometry.
 

Return to: Exp: NASA Juncture Flow - Intro Page for Transitional Symmetric Wing

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Responsible NASA Official: Ethan Vogel
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Last Updated: 04/07/2023