Reference Value for density in Fluent
 

HI

if we have Mach number and Reynolds number for an experiment then how to set Reference Value for density in Fluent to get same Cl Cd and Cm, coz if we change this value the coefficients will change considerably.

so any ideas ????????

Hope i get u correctly.

based on formula u can fix the density.
Re = rho * V * D / viscosity.

If you have access to the GUI, it is on the Reference Values pane under Problem Setup. This is where all the reference values are set.

I do not remember the TUI commands, but if you must use TUI then I'm sure you can navigate your way to find it.

Re = rho * V * D / viscosity.

ok in this relation what should be viscosity then???? like i take example of RAE 2822 airfoil, there are two links i would like to share

http://www.cfd-online.com/Forums/cd-...imulation.html

http://my.fit.edu/itresources/manual...p_VMFL017.html

in link 1 of the same CFD forum we get Density: 0.469604kg/m^3 Static Pressure: 40449.81Pa and in link 2 we get density=0.5082372kg/m^3 Static Pressure: 43765Pa.

so for same same now we compute density and we change viscosity.... and this could be done vice versa...

so my question is should we change density or viscosity our self to get same coefficients????? coz for most of the cases we dont have experimental data or reference data so what practice should we adopt

sorry, I'm having a very hard time trying to understand your question. Can you please explain it better?

To set reference values, I have explained. Now what are you having trouble with?

The definition of lift and drag coefficients as calculated by Fluent can be found in the help manual.

Are you trying to set reference values or are you trying to change fluid properties? Why would you want to change reference values to match the same Cd? You should be controlling Reynolds number or Mach number by adjusting your working fluid properties or operating conditions. Also, please understand that there is a difference between the fluid property called density and viscosity, and the reference values called density and viscosity.

Changing your reference density and viscosity does not affect the solver in any way. It only changes the computed values (Cd and such) that uses those references.

I'm sorry, if you can ask your question more clearly I can maybe answer it better. You can freely edit your posts so there is no excuse to leave it hard to understand.

ok ill try to ask more clearly

I am trying to carryout CFD analysis of RAE 2822 airfoil case 9 of agard report 138....

case 9 has Mach number of 0.73 and reynolds number of 6.5e6....

now in fluent in am using pressure farfield boundar conditon in which i have to input pressure and mach... for mach there is no problem but problems come for pressure....
should i use STP i.e 101325 Pa and 288.15 K and standard viscosity 1.7894e-5?????
or should i calculate the pressure using density from reynolds number relation.

at the end i will be comparing lift and drag coefficients obtained from experiment....
now comes the reference values so what should be my ref values to get similer cl and cd..

i am sorry for taking your time but ill be thnkful if you help me in this regard

Okay, I'm starting to understand this much better.

if you want to compare directly to the Cl and Cd they used, then use the same definition for your reference values. That means use the same type of reference values (not necessarily the same numbers). i.e. if they used inlet density then you also use inlet density (although the actual numbers may be different).

the farfield pressure should be set to the far field pressure. If you do not know what pressure to set this to, then how do you know you have a far field pressure boundary? The inlet pressure should be the inlet pressure (stagnation pressure in this case). Depending on what type of boundary condition you are using at the inlet, the pressure may be iterated.


I went back to read these two references more thoroughly. They do not define drag coefficient anywhere. This will put you in trouble matching their results since you do not know what reference values they used.

i am still in some confusion !!!!!
in the second reference
http://my.fit.edu/itresources/manual...p_VMFL017.html
the cd value is 0.0168 (exp) and 0.0165 (fluent) but to obtain this close value we have to alter something so what shoudl be that something .... and in this case we know this value of cd but as i asked earlier that if we dont know any reference value then how we will obtain true or close to true cd value???????

earlier you asked for first, how to set reference values. I answered that clearly. Then you asked for how to calculate/set pressure, I also answered that clearly. I cannot read your mind, we are interacting through these words on this forum, it really is your fault for not being able to ask the right questions clearly.

Backtrack their work and try to figure it out. The 2nd link also referenced P.H. Cook, M.A. McDonald, and M.C.P. Firmin. “AEROFOIL RAE 2822 Pressure Distribution and Boundary Layer and Wake Measurements”. AGARD Advisory Report. No. 138. 1979. It is not that bad! The reference area/length/density/viscosity and whatnot will be chosen at convenient values! The reason for using these drag coefficients was so that they could be conveniently calculated in the first place. I don't think it will be that hard to just read, but read! I'm willing to bet that all the reference properties are taken from the inlet.

If you have time you can re-run their simulation exactly and then figure out what reference values you need. Since they did not provide actual values of force or pressure distributions, you cannot even recalculate the coefficients even if you wanted to without running the simulation again. These are pretty much your only two options if you want to match their Cd exactly.