Standard Cirrus Simulated Airfoil Wind Tunnel Results

Wing Airfoil Wind Tunnel Simulations

Following are wind tunnel simulations on the Standard Cirrus wing, 53 inches from the root, where the wing chord is 32 inches. The airfoil at that point is a 32% transition between Wortmann FX S 02-196 (root rib) and Wortmann FX 66-17 A II-182 (inboard end of aileron to tip). A wing loading of 6.6 lb/sq ft was assumed.

That station was chosen because drag testing was performed on Standard Cirrus #60 at that wing station to measure the effectiveness of various drag reduction techniques.

Since the wing is a complicated three dimensional shape with changing chord, airfoil, and twist, a simple 2D analysis cannot give accurate results. For these simulations, it was assumed that the wing had constant airfoil and chord, and zero twist. The actual chord length was used to get Reynolds numbers for an altitude of 5000 ft in a standard atmosphere. The errors become apparent at airspeeds below 50 kts where the simulation could not produce coefficients of lift (C_L) large enough to make the glider fly. So these simulations below should be taken as ballpark approximations of how the wing actually performs at the 53 inch span station.

The simulations were run using DesignFOIL from DreeseCode Software. The images are pressure coefficient graphs with boundary layer information. Following is an explanation, from the DesignFOIL manual, of how to interpret the graphs:

Pressure Coefficients
The Pressure Coefficient tells the engineer how the air is acting as it travels around an airfoil. When the air slams into the leading edge of the airfoil, its speed goes to zero. This makes the pressure coefficient equal to 1. As the air moves along the sides of the airfoil and speeds up, the Pressure Coefficient becomes more and more negative. A really negative Pressure Coefficient corresponds to low pressure, or suction. Most airfoils have the most negative Pressure Coefficient on the upper surface. That is why an airfoil is sucked up into the airflow. ...

Note how the Y-axis is flipped and how the negative values are on the upper side and the positive are on the lower. This is standard practice among aeronautical engineers. The solid blue line refers to the pressure coefficients on the upper surface of the airfoil. The dashed blue line refers to the pressure coefficients on the lower surface of the airfoil.

The pressure coefficient plot represents the data used to obtain those important three numbers: the Lift Coefficient, Drag Coefficient, and the Moment Coefficient.

Boundary Layer Graphics
The black dot refers to the stagnation point on the airfoil where the air is completely stopped, or stagnated. This marks the point where the boundary layer starts growing along the upper and lower surfaces. The three additional circles shown on each surface have special meanings, designated by their size.

The first circle marks the end of the laminar boundary layer and the start of the transition region.
The second circle marks the end of the transition region and the beginning of the turbulent boundary layer.
The third and final circle designates where the boundary layer MIGHT separate from the surface.

Click for Oil Flow Tests.

  Airspeed = 50 knots

  AOA = +3.3 degrees

  Re (5000 ft) = 1407416

  C_L = 0.73

  Airspeed = 55 knots

  AOA = +1.1 degrees

  Re (5000 ft) = 1542938

  C_L = 0.62

  Airspeed = 60 knots

  AOA = -0.2 degrees

  Re (5000 ft) = 1666926

  C_L = 0.54

  Airspeed = 70 knots

  AOA = -1.3 degrees

  Re (5000') = 1920670

  C_L = 0.41

  Airspeed = 80 knots

  AOA = -2.1 degrees

  Re (5000 ft) = 2217665

  C_L = 0.31

  Airspeed = 90 knots

  AOA = -2.6 degrees

  Re (5000 ft) = 2482943

  C_L = 0.25

  Airspeed = 100 knots

  AOA = -3.0 degrees

  Re (5000 ft) = 2782821

  C_L = 0.20

  Airspeed = 110 knots

  AOA = -3.3 degrees

  Re (5000 ft) = 3063936

  C_L = 0.16

  Airspeed = 120 knots

  AOA = -3.5 degrees

  Re (5000 ft) = 3382919

  C_L = 0.14