# Negative Lift Coefficients - Help

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 February 28, 2013, 14:16 Negative Lift Coefficients - Help #1 New Member   dave Join Date: Feb 2013 Posts: 3 Rep Power: 4 Hi, I am running a case for a range of NACA airfoils to see how close my CFD comes to matching the results obtained by those before hand. At the moment I am just running the analysis at the following conditions; Angle of attack = 0 degrees Chord length = 0.001 m Flow Velocity = 200 m/s Pressure = 28 000 Pa Reynold's = 1,000,000 Kinematic Viscosity = 0.0000002 m2/s This was using airfoils created using JavaFoil with a total of 121 points, connected by a polyspline. So far my results show this, with my results first and those obtained from a source second. Experimental Data at 0 degrees angle of attack, Re=1,000,000 NACA 2410 @ Re=1,000,000 Cd = 0.03836 Cl = -0.2527 Cm = -0.0777 NACA 4424 @ Re=1,000,000 Cd = 0.1248 Cl = -0.0938 Cm = 0.0544 NACA 1408 @ Re=1,000,000 Cd = 0.02795 Cl = -0.05696 Cm = -0.0002037 Existing Data at 0 degrees angle of attack, Re=1,000,000 - Roughly taken from http://airfoiltools.com/airfoil/naca4digit NACA 2410 @ Re=1,000,000 Cd = 0.01 Cl = 0.25 Cm = -0.055 NACA 4424 @ Re=1,000,000 Cd = 0.01 Cl = 0.2 Cm = -0.03 NACA 1408 @ Re=1,000,000 Cd = 0.005 Cl = 0.1 Cm = -0.03 What my main concern is that my lift coefficients are negative, I am positive the axes are the correct way round, (positive y is upwards). I am using simpleFoam as a solver, and the mesh was created in OpenFoam using the blockMesh. It would be great if someone could shed some light as to what I am doing wrong. Thank you, Dave

May 28, 2014, 15:42
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xin
Join Date: Dec 2013
Posts: 18
Rep Power: 3
Quote:
 Originally Posted by davemanson Hi, I am running a case for a range of NACA airfoils to see how close my CFD comes to matching the results obtained by those before hand. At the moment I am just running the analysis at the following conditions; Angle of attack = 0 degrees Chord length = 0.001 m Flow Velocity = 200 m/s Pressure = 28 000 Pa Reynold's = 1,000,000 Kinematic Viscosity = 0.0000002 m2/s This was using airfoils created using JavaFoil with a total of 121 points, connected by a polyspline. So far my results show this, with my results first and those obtained from a source second. Experimental Data at 0 degrees angle of attack, Re=1,000,000 NACA 2410 @ Re=1,000,000 Cd = 0.03836 Cl = -0.2527 Cm = -0.0777 NACA 4424 @ Re=1,000,000 Cd = 0.1248 Cl = -0.0938 Cm = 0.0544 NACA 1408 @ Re=1,000,000 Cd = 0.02795 Cl = -0.05696 Cm = -0.0002037 Existing Data at 0 degrees angle of attack, Re=1,000,000 - Roughly taken from http://airfoiltools.com/airfoil/naca4digit NACA 2410 @ Re=1,000,000 Cd = 0.01 Cl = 0.25 Cm = -0.055 NACA 4424 @ Re=1,000,000 Cd = 0.01 Cl = 0.2 Cm = -0.03 NACA 1408 @ Re=1,000,000 Cd = 0.005 Cl = 0.1 Cm = -0.03 What my main concern is that my lift coefficients are negative, I am positive the axes are the correct way round, (positive y is upwards). I am using simpleFoam as a solver, and the mesh was created in OpenFoam using the blockMesh. It would be great if someone could shed some light as to what I am doing wrong. Thank you, Dave
Hi, I think your mesh should be improved.

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