Question regarding CLa Calculator

[mustang60348 2]

The CLa calculator will give you the Coefficient of Lift at 0 Alpha.

 

Since the Coefficient of Lift is determined by wind tunnel testing and flight testing for each airfoil, what does this program use to determine CLa. Does it use a standard airfoil and if so , which one.

[Fubar512 1,350]

The CLa calculator will give you the Coefficient of Lift at 0 Alpha.

 

Since the Coefficient of Lift is determined by wind tunnel testing and flight testing for each airfoil, what does this program use to determine CLa. Does it use a standard airfoil and if so , which one.

 

If you're referring to Kreelin's AeroConvert, I'd assume it to be an average of several airfoils. If you try several airfoil types in JavaFoil, you'll see that while CL0 may change appreciably among various airfoils, CLa seems to be more dependent on the wing's geometry itself.

[mustang60348 2]

If you're referring to Kreelin's AeroConvert, I'd assume it to be an average of several airfoils. If you try several airfoil types in JavaFoil, you'll see that while CL0 may change appreciably among various airfoils, CLa seems to be more dependent on the wing's geometry itself.

 

Yes, CLa, of course is the Coeffecient of Lift (at Alpha) and is determined by wind tunnel testing so I wonder how he is gettting his numbers for CLa

Edited July 26, 2007 by mustang60348

[Guest Saganuay82]

Interesting what is out there when you look for it.

 

http://www.mtu-net.ru/mosseev/cd_wing.htm

 

http://www.aa.washington.edu/courses/aa101...efficient.shtml

 

http://adg.stanford.edu/aa241/drag/ssdragcalc.html

 

http://aero.stanford.edu/WingCalc.html

[bwild 5]

If i recall it correctly from class, the Cla of an ideal infinite wing is 2Pi per radial.

 

There is a correction for a finite wing based on aspect ratio, but i can't remember it. I will see if I can find it in my books somewhere.

[mustang60348 2]

If i recall it correctly from class, the Cla of an ideal infinite wing is 2Pi per radial.

 

There is a correction for a finite wing based on aspect ratio, but i can't remember it. I will see if I can find it in my books somewhere.

 

 

That doesn't sound right because it will vary with shape of the wing. Now, unless the definition of a finite one is based on a specific shape.

 

As far as I can tell, the only way to get CLa is from wind tunnel testing and creating a lift curve vs Alpha for that airfoil. Same goes for CDa (Coeffecient of Drag)

[Fubar512 1,350]

. Same goes for CDa (Coeffecient of Drag)

 

CD is the coefficient of drag. CDa is the Coefficient of Drag as determined by alpha , and is not to be confused with CD0, which is at zero lift. For the most part, CDa is a seldom used variable.

[mustang60348 2]

CD is the coefficient of drag. CDa is the Coefficient of Drag as determined by alpha , and is not to be confused with CD0, which is at zero lift. For the most part, CDa is a seldom used variable.

 

After lots of reading, I am very confused by CLa, from my 'research' , CLa is Coeffecient of Lift, which is determined in a wind tunnel (in real life) and varies by Alpha. If this is the case, WHY in the game is there a CLa table based on Mach, since CLa doesn't change with Mach, but in fact changes by Alpha.

[Fubar512 1,350]

After lots of reading, I am very confused by CLa, from my 'research' , CLa is Coeffecient of Lift, which is determined in a wind tunnel (in real life) and varies by Alpha. If this is the case, WHY in the game is there a CLa table based on Mach, since CLa doesn't change with Mach, but in fact changes by Alpha.

 

LOL.....are you serious? Think of the statement you've just made, and you'll find the answer.

[mustang60348 2]

LOL.....are you serious? Think of the statement you've just made, and you'll find the answer.

 

 

I don't see what you are saying, everything I have read says that CLa varies on an airfoil with alpha, NOT with mach. Lift varies with speed (as well as area, density and CLa) but NOT CLa.

 

What am I missing here.

 

After reading again.

 

Perhaps this is why it varies with Mach

 

Cl = 1/2(L/S*P*V2*)

 

where

L = Lift

S = Area of the wing

P = Density

V2 = Velocity squared

 

Is this what you are referring to.

 

The confusing part is that L isn't solved for without knowing CL, which would appear to be a circular formula.

 

How can the game solve for L without knowing CL and therefore you can't solve for CL.

Edited July 26, 2007 by mustang60348

[Fubar512 1,350]

CL was calculated when you introduced the required values into the application. The output was a set CLa value at the Mach value you've inputed. That is what the game needs to read in order to interpret the data.

 

Lift varies with Mach level. Do you understand now?

[mustang60348 2]

CL was calculated when you introduced the required values into the application. The output was a set CLa value at the Mach value you've inputed. That is what the game needs to read in order to interpret the data.

 

Lift varies with Mach level. Do you understand now?

 

 

Which values are you referring to when you say "when you introduced the required values into the application".

 

This is what is confusing me about the AeroConv application: When you enter CLa calcul, you enter Surface, Span, Ctip, Croot, LE Angle and Mach.

 

Now CLa, is the Coeffecient of Lift , which should need Lift, V2, S and P (density). Yet you never enter, Density or Lift into that application. So how does it give you CLa.

[bwild 5]

Just to make it more confusing, CL is a function of alpha, Mach and Reymolds number.

 

Cl in itself is a dimensionless number, but windtunnel results for the lift of a certain airfoil can only be compared to flight test results if Mach number and Reynolds number are the same, and this is why flight test still exist.

[mustang60348 2]

Just to make it more confusing, CL is a function of alpha, Mach and Reymolds number.

 

Cl in itself is a dimensionless number, but windtunnel results for the lift of a certain airfoil can only be compared to flight test results if Mach number and Reynolds number are the same, and this is why flight test still exist.

 

 

I think the main confusion I am having is that I define CL as part of an airfoil, vice part of a wing. IOW, an airfoil does not have a surface area, chord, etc. And I was expecting to take the CL value from a Lift Curve graph and input that into the game.

[Fubar512 1,350]

Which values are you referring to when you say "when you introduced the required values into the application".

 

This is what is confusing me about the AeroConv application: When you enter CLa calcul, you enter Surface, Span, Ctip, Croot, LE Angle and Mach.

 

Now CLa, is the Coeffecient of Lift , which should need Lift, V2, S and P (density). Yet you never enter, Density or Lift into that application. So how does it give you CLa.

 

 

Density is more or less a given. That one's easy to figure out

 

As I stated before, you've inputted geometry data into the required fields to help calculate a constant, which is represented by CLa. Since lift (represented CL), is effected by mach level, we have tables that tell the flight engine what multiplication factor to use at a set mach level.

[mustang60348 2]

Density is more or less a given. That one's easy to figure out

 

As I stated before, you've inputted geometry data into the required fields to help calculate a constant, which is represented by CLa. Since lift (represented CL), is effected by mach level, we have tables that tell the flight engine what multiplication factor to use at a set mach level.

 

OK, perhaps this is where i am getting confused.

 

is CL = Lift or is CL = Coefficient of Lift.

[bwild 5]

Cl is coefficient if lift, lift is a function of Cl density speed and wing area.

Edited July 26, 2007 by bwild

[mustang60348 2]

Cl is coefficient if lift, lift is a function of Cl density speed and wing area.

 

Good, that makes sense.

 

From what I gather then, this game doesn't take alpha into account when using and figuring out CL.

[Fubar512 1,350]

I think the main confusion I am having is that I define CL as part of an airfoil, vice part of a wing. IOW, an airfoil does not have a surface area, chord, etc. And I was expecting to take the CL value from a Lift Curve graph and input that into the game.

 

And that's where you're stuck. As I explained above, to the best of my knowledge, that's not what the SF flight engine "wants" to see.

 

You simply have to stop and think. Is Lift effected by velocity (Mach level)? Is it effected by angle of attack? Is it effected by the wing area, sweep angle, tip chord, and root chord (in other words, by wing geometry)?

[Fubar512 1,350]

Good, that makes sense.

 

From what I gather then, this game doesn't take alpha into account when using and figuring out CL.

 

Actually, it does.

[bwild 5]

The game calculates CL with alpha and Mach number, and then lift with:

 

L=Cl*1/2*rho*V^2*S

[mustang60348 2]

Actually, it does.

 

Does it use a formula 2piR to figure out the value for CL OR do 'we' create an alpha table for CL.

[+Dave 2,322]

Does it use a formula 2piR to figure out the value for CL OR do 'we' create an alpha table for CL.

 

My advice....let Fubar do your FM's. It will save you a lot of headache.

[mustang60348 2]

My advice....let Fubar do your FM's. It will save you a lot of headache.

 

 

I wish I could, but unfortunately I can't

[Guest Saganuay82]

Let the man work, at least we can have more FM guys.