The untwisted wing incorporated a full-span, leading-edge Krueger flap and a full-span, single-slotted trailing-edge flap. This report contains pressure data which document effects of wing configuration and free-stream conditions on wing pressure distributions. Pressure Distributions from Subsonic Tests of a NACA 0012 Semispan Wing Model An unswept, semispan wing model incorporating a NACA 0012 airfoil section was tested in the Langley 14- by 22-Foot Subsonic Tunnel. Angle of Attack and Pressure Distribution. 40ģ.4.2 Implications of Moment Coefficient Behavior. 40ģ.4.1 Moment Coefficient Behavior in Linear and Stalled Regions. 38ģ.3.3 Implications of Lift Coefficient Behavior. 37ģ.3.1 Effects of Tripped Boundary Layer on Lift Coefficient. 35ģ.2.4 Implications of Stall and Transition. 33ģ.2.3 Comparison of Pressure Coefficients to XFOIL. 31ģ.2.2 Airfoil Behavior with Turbulence Tape. 24Ģ.3 Experimental Setup and Data Collection. 7ġ.1.2 Viscous Flow in the Boundary Layer. The airfoil exhibited a leading edge stall for both laminar and turbulentĪbstract. 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 Exercise 3: Pressure Distribution on an Airfoil (Version 2).The airfoil exhibited a leading edge stall for both laminar and turbulent boundary layers. 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.
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