[Shamoon Jamshed]

Hi,

I am modelling a simple NACA 0012 aerofoil at zero alpha but I am getting lift force at even zero angle of attack. Why is it so? Lift coefficient is under e-06.

[Santos-Dumont]

Hi,

It's normal, most of the airfoils get a zero lift at around -2°. Check your profile's polar to find out (look for the "Theory of Wing sections" .pdf on the Internet).

[Santos-Dumont]

Correction, NACA 0012 gets 0 lift at 0°.
But if the Cl your calculate is at an order as low as 10^-6, you can consider it as zero.

[Shenren_CN]

Originally Posted by Shamoon Jamshed

Hi,

I am modelling a simple NACA 0012 aerofoil at zero alpha but I am getting lift force at even zero angle of attack. Why is it so? Lift coefficient is under e-06.

Could you post the mesh here?

It's very common to get non zero lift for zero alpha if your mesh if not symmetric.

[Shamoon Jamshed]

So you mean that its all because of mesh! Guys I told you that my problem is caused by the Force not the coefficient. I also get the same for a problem of cone cylinder assembly whose mesh size is 0.4 million. Still its symmetry of the grid issue? Lift force 55 N while coefficient 1e-4.

[Shenren_CN]

Originally Posted by Shamoon Jamshed

So you mean that its all because of mesh! Guys I told you that my problem is caused by the Force not the coefficient. I also get the same for a problem of cone cylinder assembly whose mesh size is 0.4 million. Still its symmetry of the grid issue? Lift force 55 N while coefficient 1e-4.

All that I was saying is that it's very likely it's caused by the mesh. And it's so simple to test if it's true. Simple take a symmetric mesh and see what number you get. Force or coefficient, are they the same thing?...

And you problem is not CAUSED by the force; instead, the "wrong" force is telling you there might be a problem.

[Shamoon Jamshed]

I am sorry I didnt get your point. What do you mean by the wrong Force? For Normal force (or Lift at 0 alpha) coefficient I give the vector 0 1 0 in Fluent.

[Shamoon Jamshed]

For cone cylinder I made a domain and then rotated it 360 to get the 3D blocks.

[hhh]

Dear friends,
In fluent, For analysis of 2D naca0012 airfoil velocity inlet we give vcostheta in x component & vsintheta in y component, v is inlet velocity.

1)here i start analysis of 3D wing, for that i take naca 0012 airfoil & extrude 100mm, in gambit and analysis in fluent i have doubt how to set velocity inlet, here i consider my velocity is 50m/s, then what is xyz component,please let me know.

2)if you consider 3D wing what is z component(vtantheta for z component) is correct or wrong, please let me know.

2) how to find angle of attack for 3D wing? please let me know

[Shamoon Jamshed]

In 3D even if you had taken the velocity inlet as the front end boundary condition you will need to specify 2D components of velocity i.e. x and y components. The third component in z direction would be zero. However, if you are certain that there is some third component and the velocity is not uniform you must need to provide data for all the three components of velocity at inlet through Define---> Profiles.

[hhh]

Dear Shamoon Jamshed,
thanks for your kind response, i take 2D naca0012 airfoil & and make 3D wing for that i extrude 100mm and meshed in gambit.my boundary conditions are velocity inlet,pressure outlet,remaining all are symmetry, and the wall(stationary) is wing. In fluent how i give all the parameters, to find lift and drag for my 3D wing, my velocity is 10m/s.


if u don't mind, please give the details step by step in fluent& please let me know.
step1--define-solver
step2--define--BC ......like that, i hope its only 10-15 steps up to iteration,

[Shamoon Jamshed]

Quite surprising that you are asking each and every bit although you know 2D. Anyway, the procedure is no different than 3D. Just define the velocity inlet all over the far-field domain (vel inlet=pr far field in case of supersonic flow). Now define the x and y components ( I assume the wing span in x z plane) of the velocity in velocity inlet bc. Select magnitude and components. Go to solve ---> monitors ---> Drag and solve ---> monitors ---> Lift and click on write and plot so that you can view the results of Cd and Cl whether converged or not. Before this set reference values of area as unit chord length(or if you have any specific value from experiment). Then run the iterations.

Originally Posted by hhh

Dear Shamoon Jamshed,
thanks for your kind response, i take 2D naca0012 airfoil & and make 3D wing for that i extrude 100mm and meshed in gambit.my boundary conditions are velocity inlet,pressure outlet,remaining all are symmetry, and the wall(stationary) is wing. In fluent how i give all the parameters, to find lift and drag for my 3D wing, my velocity is 10m/s.


if u don't mind, please give the details step by step in fluent& please let me know.
step1--define-solver
step2--define--BC ......like that, i hope its only 10-15 steps up to iteration,

[hhh]

Dear Shamoon Jamshed,

thanks for your information.

[p.bhushan2727@gmail.com]

Dear
Shamoon Jamshed,
i am doing aerofoil simultion in fluent using aateched tutorail but my soln is not conversed...in that.whatthe problem i dont get...even 50 thosan itaration has over....sry i cant able attached ithe tutorail....will u please find it from google as aeofoil tutorail

[Shamoon Jamshed]

Originally Posted by p.bhushan2727@gmail.com

Dear
Shamoon Jamshed,
i am doing aerofoil simultion in fluent using aateched tutorail but my soln is not conversed...in that.whatthe problem i dont get...even 50 thosan itaration has over....sry i cant able attached ithe tutorail....will u please find it from google as aeofoil tutorail

Lot of tutorials can be searched on google. !!!!!!!!!!!!!!!!!!