Near zero below stall as expected for a symmetrical airfoil, but rapidly became negative after stall forĮxperimental and empirical data. Pitching moment coefficient about the quarter chord remained These data were 10% lower than the empirical airfoil data found in Theory of Wing The maximum lift coefficient for the clean airfoil was 0.9 at 10 degrees angle ofĪttack, and tripped airfoil reached a maximum lift coefficient of 1.03 at 12 degrees angle of attack, aġ4% increase. Providing a beneficial test article for contrast between smooth and laminar boundary layer behavior at Experimental results showed a suction peak at less than 1% of chord, TapeĪdded to the leading edge tripped the boundary layer, and pressure distributions were taken at 8, 10, andġ2 degrees angle of attack. The airfoil was tested in a clean configuration at angles of attack of 0, 5, 8, 10, and 12 degrees. Pitching moment and the behavior of stall for laminar and turbulent boundary layers in the USNAĬlosed-Circuit Wing Tunnel (CCWT) with an NACA 65-012 airfoil at a Reynolds number of 1,000,000. The team conducted the experiment to determine the effects of pressure distribution on lift and Midshipman First Class, Aerospace Engineering Department, EA303 Laboratory report The airfoil exhibited a leading edge stall for both laminar and turbulent boundary layers.Įxercise 3: Pressure Distribution on an Airfoil (Version 2)
Pitching moment coefficient about the quarter chord remained near zero below stall as expected for a symmetrical airfoil, but rapidly became negative after stall for experimental and empirical data.
These data were 10% lower than the empirical airfoil data found in Theory of Wing Sections from Abbott and von Doenhoff. The maximum lift coefficient for the clean airfoil was 0.9 at 10 degrees angle of attack, and tripped airfoil reached a maximum lift coefficient of 1.03 at 12 degrees angle of attack, a 14% increase. Experimental results showed a suction peak at less than 1% of chord, providing a beneficial test article for contrast between smooth and laminar boundary layer behavior at the stall condition. Tape added to the leading edge tripped the boundary layer, and pressure distributions were taken at 8, 10, and 12 degrees angle of attack. The team conducted the experiment to determine the effects of pressure distribution on lift and pitching moment and the behavior of stall for laminar and turbulent boundary layers in the USNA Closed-Circuit Wing Tunnel (CCWT) with an NACA 65-012 airfoil at a Reynolds number of 1,000,000.