WingCFD performs aerodynamic analyses of two-dimensional (2-D) and finite aspect ratio (AR) airfoil sections for incompressible and compressible flow. WingCFD uses vortex lift panels to compute CD, CL and Cm,c/4 for airfoil sections using NACA four digit airfoils, streamlined, flat plate, double wedge (D'Wedge) and imported custom shapes for a wide range of 2-D and finite AR airfoils. Several NACA five-digit airfoils from Appendix III in the book Theory of Wing Sections allow the user to rapidly specify complex imported shapes. Other useful input variables include wing Reynolds number (Re) and angle of attack in degrees. Also, WingCFD produces filled color contour plots and line color contour plots for pressure coefficient, Cp and U/U0 where the number of contour levels can be specified from 3 to 256 levels. In addition, the following standard plots are produced, Cp verses chord length and U/U0 verses chord length for the upper and lower airfoil surfaces. Also, CL verses AOA, CD verses AOA, CD verses CL, CL/CD verses AOA and Cm verses AOA are quickly plotted. Finally, the total number of 2-D vortex panels that define the upper and lower surfaces of an airfoil can be specified as 100, 200 or 300. WingCFD predicted lift slope (CNa) may be specified as either incompressible or compressible for Mach number less than 1 and then inserted into the main screen flutter velocity and divergence velocity analysis.
Figure-1, WingCFD is designed to determine CD, CL and Cm for airfoils generated using the FinSim, Fin Geometry screen
WingCFD uses 2D vortex lift panels
to determine drag coefficient (CD), lift coefficient (CL) and moment coefficient
(Cm,c/4) of airfoil sections. Airfoil section shapes are automatically inserted
from the Fin Geometry screen OR specified using NACA four digit series, Streamlined, Flat
Plate, D'Wedge or Imported shape options. The following steps outline the
basic procedure used to operate WingCFD.
After defining a complex fin or wing using
this procedure click Insert CLa to incorporate lift slope (CLa)
into the main screen flutter velocity analysis. Specify other fin cross-sections
by using the pull down menu Airfoil Shapes and select either 1)
Streamlined, 2) Flat Plate, 3) D'Wedge, 4) NACA Four-Digit series or 5)
import a previously saved user-specified fin cross-section. Use
the following format to define fin geometry for each imported
cross-sectional shape where station (X) and ordinate (Y) are given in percent
of airfoil chord where LE refers to the fin Leading Edge along average chord: [Upper Surface X-location from LE],
[Upper Surface Y-location from Chord Line], [Lower
Surface X-location from LE], [Lower Surface Y-location from Chord
for each station from the LE to Trailing Edge (TE). Please be sure to save
each fin geometry file using the .txt format. Also note the
NACA TN 4197 Flutter Method does not use the lift slope (CLa)
generated by this procedure.
NACA 0012 AIRFOIL VALIDATION
Figure 2, NACA 0012 CL verses AOA and CD verses CL.
(2) NACA 63-212 AIRFOIL VALIDATION
Figure 3, NACA 63-212 CL verses AOA and CD verses CL