Boundary layer separation, based on pressure diagrams.
01-20-2009 | 01:48 PM
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From: Geneva, , SWITZERLAND
Boundary layer separation, based on pressure diagrams.
Hi again,
My question one is I have attached the images I got from Javafoil below. The airfoil I made is the one that is on the airplane kit I am building. Please guys can you tell me whereabouts does the boundary layer separate? Please just mark the spot on one of the attached images. I am guessing that the separation occurs where the blue pressure colour suddenly turns suddenly green on the flow field picture for the Cp, pressure coefficient. Also, when the more negative the Cp is, the less the pressure is there right? And the mroe positive it is, the stronger the pressure is? Am I right?
My second question is this. How do you figure out the needed wing area for lifting a certain weight? I have ane quation which states that S=W/(W/S)
where:
W= gross weight
S= wing area
Okay, I can agree with this formula as it boldly compares the weight of the airplane and it's wing loading. Reasonable enough- but not quite what is wanted here. That equation assumes that the wing is some flat planform/surface which must provide 30 g of lift per squre meter should it's wing ladoing be 30g/m^2. But, what if I am using an airfoil which DOESN't provide that much lift?
I guess that what I am trying to say is basically I want an equation better than the one above of the common one L=Cl times 1/2 times p times V^2 times S
Any person who is at least half a foot into the field of aerodynamics knows this formula. But it is a very simple formula which provides you with lift given by a wing with no freaking airfoil?! What I want is an equation where a set of input variables are the characteristics of the airfoil I chose to put onto my wing, so that the wing dimensions I get are the ones for a wing with MY airfoil on it. Get me? I guess that Cl is one factor that I must know from my chosen airfoil but is that enough? What about the Cp or Cdp, parasite drag coefficient. I guess that to get the perfect wing dimensions you need to winf tunnel test it, but as I dont have the luxury of one how else can I find the close to exact size for my wing???
Thanks guys. I appreciate the answers.
My question one is I have attached the images I got from Javafoil below. The airfoil I made is the one that is on the airplane kit I am building. Please guys can you tell me whereabouts does the boundary layer separate? Please just mark the spot on one of the attached images. I am guessing that the separation occurs where the blue pressure colour suddenly turns suddenly green on the flow field picture for the Cp, pressure coefficient. Also, when the more negative the Cp is, the less the pressure is there right? And the mroe positive it is, the stronger the pressure is? Am I right?
My second question is this. How do you figure out the needed wing area for lifting a certain weight? I have ane quation which states that S=W/(W/S)
where:
W= gross weight
S= wing area
Okay, I can agree with this formula as it boldly compares the weight of the airplane and it's wing loading. Reasonable enough- but not quite what is wanted here. That equation assumes that the wing is some flat planform/surface which must provide 30 g of lift per squre meter should it's wing ladoing be 30g/m^2. But, what if I am using an airfoil which DOESN't provide that much lift?
I guess that what I am trying to say is basically I want an equation better than the one above of the common one L=Cl times 1/2 times p times V^2 times S
Any person who is at least half a foot into the field of aerodynamics knows this formula. But it is a very simple formula which provides you with lift given by a wing with no freaking airfoil?! What I want is an equation where a set of input variables are the characteristics of the airfoil I chose to put onto my wing, so that the wing dimensions I get are the ones for a wing with MY airfoil on it. Get me? I guess that Cl is one factor that I must know from my chosen airfoil but is that enough? What about the Cp or Cdp, parasite drag coefficient. I guess that to get the perfect wing dimensions you need to winf tunnel test it, but as I dont have the luxury of one how else can I find the close to exact size for my wing???
Thanks guys. I appreciate the answers.
01-20-2009 | 02:15 PM
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From: Jonkoping, SWEDEN
RE: Boundary layer separation, based on pressure diagrams.
ORIGINAL: Lafayette
I am guessing that the separation occurs where the blue pressure colour suddenly turns suddenly green on the flow field picture for the Cp, pressure coefficient. Also, when the more negative the Cp is, the less the pressure is there right? And the mroe positive it is, the stronger the pressure is? Am I right?
I am guessing that the separation occurs where the blue pressure colour suddenly turns suddenly green on the flow field picture for the Cp, pressure coefficient. Also, when the more negative the Cp is, the less the pressure is there right? And the mroe positive it is, the stronger the pressure is? Am I right?
It was Ludwig Prandtl who explained why this is so: Because the velocity in the boundary layer decreases towards the surface of the airfoil, the kinetic energy of the fluid inside the boundary layer is less than that at the outer edge of the boundary layer. This means that while the pressure rise in the outer flow may be quite significant, the fluid particles inside the boundary layer may not be able to overcome even a small adverse pressure gradient causing fluid particles near the surface to stop and maybe even reversing the flow causing a recirculating flow region characteristic of separated flows.
As for your lift equation questions the formula L=0.5*rho'v^2*S*Cl is commonly used. What lift coefficient a certain airfoil can produce at a certain angle of attack has to be determined from polar curves and/or CFD calculations. There is no simple formula for this, especially not at high angle of attacks. At small angle of attacks the Cl vs. alpha curve has a slope of 2*pi for 2d airfoils. For finite span wings this value should be corrected for aspect ratio (and other factors). A commonly used correction for aspect ratio is 2pi*AR/(1.25+AR). This is often accurate enough for model aircraft purposes.
01-20-2009 | 02:44 PM
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From: Bloomington, MN,
RE: Boundary layer separation, based on pressure diagrams.
ORIGINAL: Lafayette
Hi again,
My question one is I have attached the images I got from Javafoil below. The airfoil I made is the one that is on the airplane kit I am building. Please guys can you tell me whereabouts does the boundary layer separate? Please just mark the spot on one of the attached images. I am guessing that the separation occurs where the blue pressure colour suddenly turns suddenly green on the flow field picture for the Cp, pressure coefficient. Also, when the more negative the Cp is, the less the pressure is there right? And the mroe positive it is, the stronger the pressure is? Am I right?
My second question is this. How do you figure out the needed wing area for lifting a certain weight? I have ane quation which states that S=W/(W/S)
where:
W= gross weight
S= wing area
Okay, I can agree with this formula as it boldly compares the weight of the airplane and it's wing loading. Reasonable enough- but not quite what is wanted here. That equation assumes that the wing is some flat planform/surface which must provide 30 g of lift per squre meter should it's wing ladoing be 30g/m^2. But, what if I am using an airfoil which DOESN't provide that much lift?
I guess that what I am trying to say is basically I want an equation better than the one above of the common one L=Cl times 1/2 times p times V^2 times S
Any person who is at least half a foot into the field of aerodynamics knows this formula. But it is a very simple formula which provides you with lift given by a wing with no freaking airfoil?! What I want is an equation where a set of input variables are the characteristics of the airfoil I chose to put onto my wing, so that the wing dimensions I get are the ones for a wing with MY airfoil on it. Get me? I guess that Cl is one factor that I must know from my chosen airfoil but is that enough? What about the Cp or Cdp, parasite drag coefficient. I guess that to get the perfect wing dimensions you need to winf tunnel test it, but as I dont have the luxury of one how else can I find the close to exact size for my wing???
Thanks guys. I appreciate the answers.
Hi again,
My question one is I have attached the images I got from Javafoil below. The airfoil I made is the one that is on the airplane kit I am building. Please guys can you tell me whereabouts does the boundary layer separate? Please just mark the spot on one of the attached images. I am guessing that the separation occurs where the blue pressure colour suddenly turns suddenly green on the flow field picture for the Cp, pressure coefficient. Also, when the more negative the Cp is, the less the pressure is there right? And the mroe positive it is, the stronger the pressure is? Am I right?
My second question is this. How do you figure out the needed wing area for lifting a certain weight? I have ane quation which states that S=W/(W/S)
where:
W= gross weight
S= wing area
Okay, I can agree with this formula as it boldly compares the weight of the airplane and it's wing loading. Reasonable enough- but not quite what is wanted here. That equation assumes that the wing is some flat planform/surface which must provide 30 g of lift per squre meter should it's wing ladoing be 30g/m^2. But, what if I am using an airfoil which DOESN't provide that much lift?
I guess that what I am trying to say is basically I want an equation better than the one above of the common one L=Cl times 1/2 times p times V^2 times S
Any person who is at least half a foot into the field of aerodynamics knows this formula. But it is a very simple formula which provides you with lift given by a wing with no freaking airfoil?! What I want is an equation where a set of input variables are the characteristics of the airfoil I chose to put onto my wing, so that the wing dimensions I get are the ones for a wing with MY airfoil on it. Get me? I guess that Cl is one factor that I must know from my chosen airfoil but is that enough? What about the Cp or Cdp, parasite drag coefficient. I guess that to get the perfect wing dimensions you need to winf tunnel test it, but as I dont have the luxury of one how else can I find the close to exact size for my wing???
Thanks guys. I appreciate the answers.
The wing area is often chosen to allow the plane to land at a desired speed. If you know the landing speed you would like, and the Cl that the wing can generate, then you can calculate the area required. I don't know what the maximum Cl is for a NACA 3412, but at the speed a model would fly, I am guessing it would be around 1.0. Pick your desired landing speed, assume Cl=1.0, and calculate the resulting area.
banktoturn
01-20-2009 | 02:57 PM
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From: Geneva, , SWITZERLAND
RE: Boundary layer separation, based on pressure diagrams.
Yeah the lift equation you wrote is the one I mentioned in my post. I just put p for rho as that is the notation for freestream density the book I have uses; it has a rotated by a right angle 8 as well, showing it is freestream density.
I understand all you said except how have you pointed out that in fig. 1 (i.e. the graph) the separation ocurs at 30% of the way back from the LE? I would this that it is at 0.2 (i.e. 20%) because that is where the pressure rises. Can you reply on this please? thanks.
Lafayette.
thanks for responding to my questions!!!
I understand all you said except how have you pointed out that in fig. 1 (i.e. the graph) the separation ocurs at 30% of the way back from the LE? I would this that it is at 0.2 (i.e. 20%) because that is where the pressure rises. Can you reply on this please? thanks.
Lafayette.
thanks for responding to my questions!!!
01-20-2009 | 03:06 PM
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From: Bloomington, MN,
RE: Boundary layer separation, based on pressure diagrams.
ORIGINAL: Lafayette
Yeah the lift equation you wrote is the one I mentioned in my post. I just put p for rho as that is the notation for freestream density the book I have uses; it has a rotated by a right angle 8 as well, showing it is freestream density.
I understand all you said except how have you pointed out that in fig. 1 (i.e. the graph) the separation ocurs at 30% of the way back from the LE? I would this that it is at 0.2 (i.e. 20%) because that is where the pressure rises. Can you reply on this please? thanks.
Lafayette. thanks for responding to my questions!!!
Yeah the lift equation you wrote is the one I mentioned in my post. I just put p for rho as that is the notation for freestream density the book I have uses; it has a rotated by a right angle 8 as well, showing it is freestream density.
I understand all you said except how have you pointed out that in fig. 1 (i.e. the graph) the separation ocurs at 30% of the way back from the LE? I would this that it is at 0.2 (i.e. 20%) because that is where the pressure rises. Can you reply on this please? thanks.
Lafayette.
thanks for responding to my questions!!!
suspicious to me. It occurs before the thickest point on the wing, which
doesn't seem right, and it is very abrupt. I am viewing that as a likely
spurious value. If JavaFoil gives you the option to run a more precise analysis,
I would be tempted to do that.
Boundary layer separation, if it occurs, will tend to occur a short distance
after the negative pressure peak.
banktuturn
01-20-2009 | 03:33 PM
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From: Jonkoping, SWEDEN
RE: Boundary layer separation, based on pressure diagrams.
ORIGINAL: Lafayette
I understand all you said except how have you pointed out that in fig. 1 (i.e. the graph) the separation ocurs at 30% of the way back from the LE? I would this that it is at 0.2 (i.e. 20%) because that is where the pressure rises. Can you reply on this please? thanks.
I understand all you said except how have you pointed out that in fig. 1 (i.e. the graph) the separation ocurs at 30% of the way back from the LE? I would this that it is at 0.2 (i.e. 20%) because that is where the pressure rises. Can you reply on this please? thanks.
A short comment: What you should look for is the chordwise position where the slope of the pressure distribution becomes negative. This is where the pressure gradient becomes adverse. What you consider to be a pressure rise at slightly less than 20% is actually a pressure decrease (look at the sign of Cp).
