[Venkatraman N]

Hello.I am trying to run a compressible supersonic flow simulation in ansys fluent.The geometry has an inlet and a pressure outlet.I am having trouble finding out what exactly I should initialize the boundary conditions with.Could someone provide links to youtube videos or articles related to compressible supersonic flow in ansys?

Thank you.

[LuckyTran]

You mean you need help with the boundary conditions? Boundary conditions are boundary conditions, they do not need initialization like the interior.

The inlet could be up either a mass-flow inlet or a pressure inlet. It's up to you, the modeller. The problem does not begin until you have defined your boundary conditions, and reading up on how Ansys works will do very little to help you. What do you want to simulate exactly?

[Venkatraman Nagarajan]

I am trying to simulate a compressible supersonic flow along a truncated aerospike nozzle.I understand that I need to set the inlet as a pressure inlet and the outlet as a pressure outlet.In the pressure inlet,I need to specify values for Total gauge pressure and supersonic/initial gauge pressure.I need to know what these quantities represent.When i give the gauge pressure to be 70bar(according to our design) and the supersonic/gauge pressure to some non zero value (say 10 pa) I am experiencing divergence.I set the density to be that for an ideal gas.I set the operating conditions to 0 pascal and the pressure outlet to be 101325pa.

[LuckyTran]

Is your inlet supersonic? The supersonic gauge pressure is only needed if the inlet is supersonic. You should put the static gauge pressure here. Since you have specified the operating pressure to be 0, you can of course use absolute values.

A combination of 70bar and 10Pa would be a Mach number of practically infinity.

If your inlet is not supersonic, then something else is a problem. Make sure you have a really good initial guess. Since density is calculated from the pressure and temperature, your density field on the first iteration will be way off and divergence is really common. I find it helpful to freeze the energy equation for the first 30-100 iterations and let the pressure & density fields become more realistic. Then I unfreeze the energy equation, and sometimes reduce under-relaxation factors to let the temperature field develop.

Btw you can also use a mass-flow rate inlet, it's similar to the stagnation pressure inlet. And don't forget your inlet stagnation temperature!

[Venkatraman Nagarajan]

Yes my inlet is supersonic.I obtained the values of pressure at chamber,throat and the exit from the NASA CEARUN website.I am not sure which values to use for the initial guess at the boundary conditions,that is the 'Total gauge pressure' and 'supersonic/inlet gauge pressure' at my pressure inlet in Fluent setup.Is there a way I could determine these two?

Also I am using 101325pa(1atm) at my pressure exit because I want the flow to expand to atmospheric conditions.I am using a K-epsilon model,because my main idea is to observe eddies at the base of the 'aerospike' and provide suitable base bleed.

[LuckyTran]

Boundary conditions are not initial guesses! They are boundary conditions!

The "inlet" stagnation pressure goes into the total pressure field.
The "inlet" static pressure goes into the supersonic pressure field. 10Pa is way too low it shouldn't be less than 101325 Pa in a truncated nozzle.

Your computational inlet of course is the corresponding section of your nozzle. I'm guessing the inlet of your aerospike is the after the throat, otherwise it can't be supersonic. The stagnation pressure is roughly the chamber pressure (minus a few percent). You can back out the static pressure from the expected Mach number or something, but it will affect your mass-flow.

I've only used NASA CEA as a chemical equilibrium solver, I'm not sure how you used it to get pressure at different locations. I don't really get what you can obtain from CEA other than an adiabatic flame temperature. So please enlighten me.

Your outlet pressure and k-epsilon make sense. Just check your supersonic pressure and initial conditions and then play with it.

[Venkatraman Nagarajan]

Ok.Thanks.I ran NASA CEA for a rocket which I learnt from one of my courses.It gives you the pressure and a lot other quantities at chamber,throat and the exit.Its very simple.Go to NASA CEA.Choose rocket.And then feed it with the values it asks for,example pc/pe,air to fuel ratio etc.You should be able to get the results.

And I will try giving the values a bit differently.If my total stagnation pressure is 70bar,what would be the value of static pressure?Is there a way to calculate it?Or do I give a value randomly?

[LuckyTran]

Ahh I see.

What pressure is the CEA output then? Isn't it the static pressure? The total pressure should be ~constant and pretty close to the chamber pressure. I think you can just take the static pressure from CEA if one of those stations corresponds to your inlet.

Alternatively, you can just calculate the static pressure from isentropic relations using the stagnation temperature and Mach number.

The supersonic static pressure is not random. The static pressure you put here will determine the inlet Mach number, which is pretty important for your case.

[Venkatraman Nagarajan]

Yes.Thank you so much sir.Your help is appreciated.I will definitely do as per your advise.Thank you once again.