[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

[guxin7005]

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.