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.

Step 1. The specification of Streamlined, Flat Plate and D'Wedge airfoil section shapes requires the specification of fin thickness in terms of maximum thickness in percent chord (Tmax/Chord X 100). Maximum fin thickness in percent chord for all airfoil types is specified by using the third data entry box on the first line, NACA Four-Digit series airfoil description. The first two data entry boxes for Streamlined, Flat Plate and D'Wedge airfoil section shapes are disabled. However, for the NACA Four-Digit series, the first two data entry boxes are enabled, where the first data entry box refers to maximum camber in percent chord and the second data entry box refers to position of the maximum camber in tenths of a chord from the leading edge (LE).

The following definitions are needed to define camber and camber location for NACA airfoils. First, the mean camber line is the locus of points halfway between the upper and lower surfaces of the airfoil as measured perpendicular to the mean camber line. Then, the chord line is a straight line that connects the leading and trailing edges of the airfoil and is simply referred to as the chord of the airfoil and is usually defined using the symbol, c. Using these definitions the camber is the maximum perpendicular distance between the mean camber line and the chord line of the airfoil. Camber location is simply located as a percentage of the chord length from the leading edge of the airfoil.

The complete specification of four-digit NACA airfoils and standard airfoils are summarized below for the first line, NACA Four-Digit series airfoil description. Where the first two spaces are disabled for Streamlined, Flat Plate, and D'Wedge airfoil section shapes but are required for NACA four-Digit airfoils. [Max camber in percent chord], [Position of max camber in 1/10th chord], [Max thickness in percent chord].

Step 2. The Reynolds number (Re = rUc/m) of the fin is defined on the second line of input data. Reynolds number is defined as the ratio of the inertial forces represented by the density of the medium (r), free stream velocity (U), and fin dimension (c) to the friction forces in the boundary layer represented by the viscosity of the medium (m). The following Reynolds number calculator is useful for computing Reynolds number for WING CFD (Note: This off-site calculator has not been checked for accuracy). For more information and theory about Reynolds number please visit the Wikipedia on-line encyclopedia.

Step 3. Fin angle of attack (a) is defined relative to the chord line for all section shapes on the third line of input data.

Step 4. Fin aspect ratio (AR = Span/Chord) is defined on the fourth line of input data. Aspect ratio must be non-zero and Checked to be included in the computation of CD, CL and Cm,c/4. The Aspect ratio input allows an "approximate solution" of end effects and 3D wings.

Step 5. Select one of five fin section shapes using the Airfoil Shapes pull-down menu. Streamlined, NACA four-digit, Flat Plate and D'Wedge section shapes are directly drawn after selection. In addition, arbitrary fin section shapes may be defined using the Import X-Y command. Many NACA Five-Digit section shapes are included in NACA_AIRFOILS.zip (located in the FinSim directory) are drawn by using the Import X-Y command after unzipping the file.

Import File format for each station (X) and ordinate (Y) given in percent of airfoil chord (LE is the Leading Edge): [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 Line] for each station from the LE to TE.

Step 6. Perform a 2D Vortex Panel aerodynamics analysis by clicking the SOLVE command button and follow the instructions displayed in the lower left status bar. Instructions displayed in the Status bar will state when a valid solution is achieved and when it is permissible to click the various plot command buttons.

Step 7. Display results using the following commands in the Plots pull-down menu: Cp verses X, U/U0 verses X, CL verses AOA, CD verses AOA, CD verses CL, CL/CD verses AOA, Cm verses AOA, U/U0 Contours (Filled and Line) and finally Cp Contours (Filled and Line). Where AOA refers to angle of attack in degrees, Cp = (P - PINF) / q = 1 - (U/U0)^2 is the pressure coefficient and U0 refers to the free stream velocity. Reference: THEORY OF WING SECTIONS, by Abbott and Doenhoff.

Figure-2, This NACA 150512 wing section illustrates the simplicity of specifying arbitrary fin cross-sectional shapes using FinSim

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 Line] 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.

NOTES:
1) Reference: THEORY OF WING SECTIONS, by Abbott and Doenhoff.
2) FinSim's WingCFD results (red dots) generated using the Save Results As command under File. The Results were plotted using Excel (or any spreadsheet program) and compared to THEORY OF WING SECTIONS data for the NACA 0012 and NACA 63-212 wing sections.