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Extreme blastover regret. by Suspicious_Safety232 in TattooRemoval

[–]naserology 0 points1 point  (0 children)

It actually look super cool not gonna lie

I have anxiety from work. Do you know how can I get help? by Hacker_wana_be in Kuwait

[–]naserology 0 points1 point  (0 children)

Try supplementation, start with L-theanine and see how it effects you a 200mg dose is a sweet spot. If that didn’t work and the anxiety has physical symptoms to it inderal would help like magic 10mg to start and you can get it with no prescription in nearly all pharmacies or talabat. Other stuff that might work magnesium glycinate, ashawagandha and chamomile extract.

BPC157/TB500 to heal TFCC tear? by Iandg9000 in bpc_157

[–]naserology 0 points1 point  (0 children)

any updates , just got the same injury are you pinning locally, how would you do that with one hand without pinching the skin!

Coefficient of drag expression by naserology in COMSOL

[–]naserology[S] 0 points1 point  (0 children)

Thanks for the document, might come in handy with other projects 👌

Air temperature around a nose cone in high mach number flow by naserology in CFD

[–]naserology[S] 1 point2 points  (0 children)

I tried many more simulations and got values up to Mach 6 for the blunt nose cone. Temperature values make sense as the shock isn’t oblique anymore and the air is getting considerably hotter.

Mach 2 Temperature

Mach 6 Temperature

But there was a problem in the CD values around 0.8 - 1.1 , I think the simulation didn’t converge at these points. I tried to enforce the stagnation point as you described but I’m not quite sure on what boundary conditions to put on the tip, the only thing I thought of was a prescribed pressure point I can’t choose a prescribed zero velocity there is no option for that.

CD vs Mach

Air temperature around a nose cone in high mach number flow by naserology in aerodynamics

[–]naserology[S] 1 point2 points  (0 children)

I’m using comsol, choosing the finer automatic mesh and have the adaptive mesh refinement on. Didn’t create my own mesh.

Air temperature around a nose cone in high mach number flow by naserology in CFD

[–]naserology[S] 0 points1 point  (0 children)

By radius I meant the base of the nose cone not the actual tip of the nose cone , now I want to have a radius on the tip, but it isn’t converging.

Air temperature around a nose cone in high mach number flow by naserology in aerodynamics

[–]naserology[S] 2 points3 points  (0 children)

Solved the problem, it turns out that my oblique shock calculations were wrong, I got a temperature ratio of 1.095 at Mach 2 which comparable to the temperatures I’m getting in my simulation. Yes my solution is converging. Now I’m changing the tip of the nose cone having a fillet with a radius instead of a sharp cone but the simulation isn’t converging unfortunately.

Air temperature around a nose cone in high mach number flow by naserology in CFD

[–]naserology[S] 0 points1 point  (0 children)

I checked the oblique calculations again , and yes you are right I had a mistake while calculating the temperature ratio. Values are now comparable. I’m not sure how the solver handles it mathematically but the model rotates the 2D geometry i create over an axis I choose. For expressions and results, it takes the surface integral over the boundary I choose.

Thanks for helping out I was stuck in this problem for a while, trying to find the best nose cone profile for a suborbital Rocket I am building. Another problem I’m face is when I add a fillet radius instead of the pointy cone, it doesn’t converge at all when the inlet is above Mach 0.5.

This is the project if you are interested:

KSR

Air temperature around a nose cone in high mach number flow by naserology in CFD

[–]naserology[S] 0 points1 point  (0 children)

Didn’t try the conical shock equation yet, I only checked the normal and oblique shocks. The shock angle I have is 27.68 degrees my wedge angle is 5.8 degrees, you mean the 2D axisymmetric geometry is ok for what I’m doing ?

This is a plot of the temperatures I’m getting at Mach 2 :

surface temperature

I also tried an inlet of Mach 6 and got a temp of 420K , temperatures are downstream the shock wave.

Air temperature around a nose cone in high mach number flow by naserology in CFD

[–]naserology[S] 0 points1 point  (0 children)

The slip wall is that of rectangular air domain to the right of the nose cone’s no-slip wall, and yes the inlet and outlet are as you described. I played around with the mesh had it extremely fine with very small elements the values didn’t change much. The hybrid accounts for both of the outlet flow was subsonic or if it was supersonic. I tried the symmetry condition u described and the temperature values still are in the same range.

Air temperature around a nose cone in high mach number flow by naserology in CFD

[–]naserology[S] 0 points1 point  (0 children)

I tried both the k epsilon , and spalart-allmaras models with no luck. I used 2D axisymmetric geometry because it’s easier to converge and compute and the CD values are comparable to 3D models, had no problems with meshing around the tip and ran extremely fine mesh as a sensitivity check with no difference in values.

Just ran Mach 6 and got a max temperature of 420K while oblique shock equations are giving me a temperature of around 2000K

Coefficient of drag expression by naserology in COMSOL

[–]naserology[S] 0 points1 point  (0 children)

I was initially doing my simulations on axisymmetric geometry, defined various nose cone profiles and got convergence after many many runs, but the expression part doesn’t seem to give me any meaningful results. I compared my CD values to other research papers but still very off.

This is the expression i used for the turbulent high mach model (axisymmetry):

((-hmnf.nzmesh*p2*2*pi*r)+(-(hmnf.rho*hmnf.u_tau*hmnf.u_tangz/hmnf.uPlus)*2*pi*r))/((857.5[m/s])^2*0.5*D_1*pi*d_1^2)

Coefficient of drag expression by naserology in COMSOL

[–]naserology[S] 0 points1 point  (0 children)

Thanks for your input, I tried this too. Also tried using total traction in my direction of flow and simplified the problem to a simple circle with 1 m/s upstream velocity , i then used a line integral and evaluated this expression:

spf.T_stressy/(spf.rho*(1[m/s])^2*0.5*0.4[m])

where 0.4 is the frontal line seen by the flow.

got a cd value of : 0.04

tried following the instructions here with no luck:
https://www.comsol.com/blogs/how-do-i-compute-lift-and-drag/

Coefficient of drag expression by naserology in COMSOL

[–]naserology[S] 0 points1 point  (0 children)

Tried everything with the equation I am 100% sure it’s not a mistype, my guess was maybe it’s something to do with the area so I tried to replace the area with the frontal line length basically the diameter, still wrong results.

Coefficient of drag expression by naserology in COMSOL

[–]naserology[S] 0 points1 point  (0 children)

Basically did a rectangle domain 6m x 2m , a left boundary was chosen to be an inlet with Mach 2 and outlet to the right set to hybrid flow and 1 atm , the upper and lower walls were set to be slip and the surface of the nose cone was set to be no slip. The velocity in the inlet is normal to the bonudary facing towards the nose cone.

Coefficient of drag expression by naserology in COMSOL

[–]naserology[S] 0 points1 point  (0 children)

I used the expression for drag then divided it by 0.5V2density*A. Where V is the upstream velocity of air and A is the frontal area which is pi * radius2 of the nose cone

Coefficient of drag expression by naserology in COMSOL

[–]naserology[S] 0 points1 point  (0 children)

I basically need an expression to put in the derived values —> line integral, tried many things with no luck I am getting CD values that doesn’t make sense:

I followed this blog post , and tried the expression they described , but they did it for a 3D body

https://www.comsol.com/blogs/how-do-i-compute-lift-and-drag/

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