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NACA0012 airfoil Critical reynolds number

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Old   December 7, 2012, 13:58
Default NACA0012 airfoil Critical reynolds number
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hi guys

i need a resource for critical Reynolds number over NACA0012 and NACA0015 and NACA0018 airfoils. I search a lot but i can't find a paper or part of book that report these numbers. can any body help me and recommend me a valid resource.

thx
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Old   December 7, 2012, 15:31
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dear it can be calculated by using hydraulic dia meter
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Old   December 7, 2012, 18:43
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Originally Posted by sajid View Post
dear it can be calculated by using hydraulic dia meter
thx, but how i can calculate the critical Reynolds umber via hydraulic diameter for airfoils?
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Old   December 7, 2012, 19:34
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If you are trying to establish the transition point, the first approximation is to use the flat plate value of Re = 500000, where the distance is measured from the nose of the airfoil (or the attachment point for high angles of attack). Googling critical Reynolds airfoil may provide you with more info.
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Old   December 7, 2012, 19:45
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Quote:
Originally Posted by agd View Post
If you are trying to establish the transition point, the first approximation is to use the flat plate value of Re = 500000, where the distance is measured from the nose of the airfoil (or the attachment point for high angles of attack). Googling critical Reynolds airfoil may provide you with more info.
i use google with different keyword but ican't find nothing proper. but i found it's diffrent value for diffrent angles of attack and a major question for me:
in angle of attack of 0 degrees we don't have a separation point on NACA 0012 airfoil (i use the fluent for simulation) at re=500,000 , it that mean the flow is laminar. but it shouldn't be correct, because over the flat plate the flow is turbulent for re>500,000 and on the airfoils we have thickness on the critical Reynolds number should be smaller.

is it correct?

otherwise in the Huang et al. work (2004), this is my base type for validation, the flow with re=500,000 was introduced fully turbulent flow over NACA0012 airfoil.
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Old   December 7, 2012, 23:58
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I'm not sure what you are saying here, but laminar flow over a NACA0012 will separate. Like I said, Re = 500000 is only a rough guide. For more info you could look here


www.risoe.dk/rispubl/VEA/veapdf/RIS-R-987.pdf

if this link works. I found it by googling "critical reynolds number airfoil"
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Old   December 8, 2012, 08:28
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Use X-foil
Attached Images
File Type: jpg Naca0012.jpg (30.1 KB, 84 views)
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Old   December 8, 2012, 08:39
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Originally Posted by truffaldino View Post
Use X-foil
thx a lot
i don't know how to use X-foil. can you explain how to run X-foil?
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Old   December 8, 2012, 09:28
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It is extremely easy to lern and use. Google x-foil tutorial or x-foil manual
or look at

Session for NACA0012 will look as follows:

opening x-foil you will get prompt XFOIL c>, you just enter naca0012 like this

XFOIL c> NACA0012 + enter

the x-foil will generate naca0012 coordinates and will print the foil description followed by XFOIL prompt, then you put

XFOIL c> OPER + enter

you will get new prompt .OPERi, it means that you entered flow analysys subroutine. Then you put

.OPERi c> visc 5E5 + enter

i.e. you fixed your reynolds number to 500K and you will get prompt OPERv

then you put

.OPERv c> alfa 0 +enter

it means your angle of attack is 0: the xfoil will show you analysys data, if it writes that computation does not converce, put again alfa 0 etc unil it converges. If you want re-initialize computations put init

To exit from analysys subroutine (or any other sub-routine) to higer level routine just press enter.

To change free stream turbulence parameters enter

.OPERv c> vpar + enter

you will get prompt VPAR. By default Ncrit=9, to decrease free stream turbulence put higer Ncrit, e.g.

..VPAR c> N 15 + enter

then just press enter to return to OPERv and put your alfa


For help press ? + enter

To see transition and separation points use command

..OPERv c> vplot + enter

you will get prompt VPLO

Then put

...VPLO c> CF + enter

you will see plot Cf vs X, the point where Cf=0 is the separation point
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Old   December 8, 2012, 10:16
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Quote:
Originally Posted by truffaldino View Post
It is extremely easy to lern and use. Google x-foil tutorial or x-foil manual
or look at

Session for NACA0012 will look as follows:

opening x-foil you will get prompt XFOIL c>, you just enter naca0012 like this

XFOIL c> NACA0012 + enter

the x-foil will generate naca0012 coordinates and will print the foil description followed by XFOIL prompt, then you put

XFOIL c> OPER + enter

you will get new prompt .OPERi, it means that you entered flow analysys subroutine. Then you put

.OPERi c> visc 5E5 + enter

i.e. you fixed your reynolds number to 500K and you will get prompt OPERv

then you put

.OPERv c> alfa 0 +enter

it means your angle of attack is 0: the xfoil will show you analysys data, if it writes that computation does not converce, put again alfa 0 etc unil it converges. If you want re-initialize computations put init

To exit from analysys subroutine (or any other sub-routine) to higer level routine just press enter.

To change free stream turbulence parameters enter

.OPERv c> vpar + enter

you will get prompt VPAR. By default Ncrit=9, to decrease free stream turbulence put higer Ncrit, e.g.

..VPAR c> N 15 + enter

then just press enter to return to OPERv and put your alfa


For help press ? + enter

To see transition and separation points use command

..OPERv c> vplot + enter

you will get prompt VPLO

Then put

...VPLO c> CF + enter

you will see plot Cf vs X, the point where Cf=0 is the separation point
thx a lot, it was very useful
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Old   December 9, 2012, 23:43
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Dear Kianoosh;
I don't want to make you confused, but it is so hard to say the exact Re in which transition starts. It depends on a lot of parameters, such as turbulence intensity, pressure gradient.
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Old   December 10, 2012, 02:15
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Quote:
Originally Posted by mb.pejvak View Post
Dear Kianoosh;
I don't want to make you confused, but it is so hard to say the exact Re in which transition starts. It depends on a lot of parameters, such as turbulence intensity, pressure gradient.
i agree with you absolutely, this is very difficult work.
i try to define th critical Re in this way. for naca0012 airfoil i starting to found the separation point from Re=5e5. in this Re number the separation point was at aboiut 0.85 C. I try with different Re Number 4e5, 3e5, 2e5, e5, 8e4 until the flow over airfoil to be laminar or separation point be in 1. this mean is we don't have any separation on the airfoil surface at Critical Re.

i the Critical Reynolds Number for NACA0012 about e5 and for NACA 0015 about 3.5e4. whats you idea?
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Old   December 10, 2012, 03:43
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Dear kianoosh;
I recommend you to read "Boundary-Layer Measurements on an Airfoil at Low Reynolds Numbers" by M. Brendel. it is very helpful paper in describing transition and how it can be found out where flow starts laminar to turbulent transition.
I hope it can answer your questions.
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Old   December 15, 2012, 10:04
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Is the paper by Brendel available online.

If yes, kindly post the link.

else kindly mail me at debanik1@gmail.com
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Old   December 15, 2012, 12:07
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Originally Posted by debanik1 View Post
Is the paper by Brendel available online.

If yes, kindly post the link.

else kindly mail me at debanik1@gmail.com
The paper available online. You can find it by googling the title of paper. If you can't find it message me to send it for you.
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Old   December 15, 2012, 12:10
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kindly mail me at debanik1@gmail.com
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Old   December 15, 2012, 12:15
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Quote:
Originally Posted by mb.pejvak View Post
Dear kianoosh;
I recommend you to read "Boundary-Layer Measurements on an Airfoil at Low Reynolds Numbers" by M. Brendel. it is very helpful paper in describing transition and how it can be found out where flow starts laminar to turbulent transition.
I hope it can answer your questions.
I read this paper , but i can't find that my way to found critical Re number is correct or no. What's your idea to my describe way above. It's important that I found critical Raynolds number In angle of attack of zero degree.
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Old   December 15, 2012, 12:18
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Quote:
Originally Posted by debanik1 View Post
kindly mail me at debanik1@gmail.com
Sure, But i'm not at home now. I send it in next 4 hours.
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Old   January 14, 2014, 10:51
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please, send to vaina74@hotmail.com.
thanks.
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