Effects of Altitude on Required Coefficient of Lift

[TequilaChaser]

rgr Stoney, thanks for the details.......I did not even notice you started this thread back nearly 2 yrs ago in 2008, until you just mentioned it.....

will keep an eye on this thread to see if I can learn something new, thx again 

[Charge]

"If you know the weight (equal to required lift) and wing area of the plane, and dynamic pressure (equal to the sum of the 1/2pV^2 expression of the lift equation), you can get the required Cl for that condition."

Ok, so you mean that weight is similar to the lift (L) for a wing at a specified angle of attack, i.e. it can be used to determine Cl. I understand that the wing-area would be meaningless if the profile lift i.e. Cl was unknown and what you want from the equation is lift (L) i.e. weight. 

http://en.wikipedia.org/wiki/Lift_(force)

http://en.wikipedia.org/wiki/Lift_coefficient

-C+

[Stoney]

Well, all your doing is taking the basic lift equation and solving for Cl (as in the Coefficient of Lift equation on the Wiki page).  If lift = weight, then you can substitute the weight of the aircraft into the equation for L (lift), determine the dynamic pressure, and use the known wing area to come up with the required coefficient of lift for that condition.

The result is the Cl for the entire aircraft, and not the wing only, or airfoil.

[Ex-jazz]

The result is the Cl for the entire aircraft, and not the wing only, or airfoil.

I think that would require a whole plane area data, not just a wing.

[Wmaker]

I think that would require a whole plane area data, not just a wing.

Well, as I understand it, the CLmax represents that data in itself. In other words, you can have two aircraft that have the same wingarea but because the other aircraft (whole aircraft, not just the wing) has higher lift coeffient it will stall at a slower speed at certain altitude. Lift coefficient is just a dimensioless figure that helps in understanding the lift characteristics of an aircraft.

[Charge]

So, if you calculate Cl (say, with alpha 15deg for certain wing-profile) in desired condition and from that you can calculate L (lift) =weight that would result in, say 10.000kg you would know that with that alpha a plane with that Cl and weighing 3000kg (at certain air pressure) could pull a bit over 3Gs in conditions that were used to calculate the Cl in the first place? Also in effect of that the equation can give you the alpha at which that 3000kg aircraft would be flying level i.e. the L= "weight" would be 3000kg?   

Calculating the Cl is tricky though as was evident in NACA document Gripen posted in other thread...

http://naca.central.cranfield.ac.uk/reports/1946/naca-tn-1044.pdf

PS. Bear with me, please, I really suck at math...   

-C+

[Stoney]

So, if you calculate Cl (say, with alpha 15deg for certain wing-profile) in desired condition and from that you can calculate L (lift) =weight that would result in, say 10.000kg you would know that with that alpha a plane with that Cl and weighing 3000kg (at certain air pressure) could pull a bit over 3Gs in conditions that were used to calculate the Cl in the first place? Also in effect of that the equation can give you the alpha at which that 3000kg aircraft would be flying level i.e. the L= "weight" would be 3000kg?   

Calculating the Cl is tricky though as was evident in NACA document Gripen posted in other thread...

http://naca.central.cranfield.ac.uk/reports/1946/naca-tn-1044.pdf

PS. Bear with me, please, I really suck at math...   

-C+

You don't need AoA at all for this calculation.  What this tells you is the required Cl for that condition.  If the plane is capable of performing that maneuver, you know it's capable of that computed Cl.  Therefore, if you know the stall speed, for example, of a plane at a certain configuration, then you can compute the Clmax the airplane is capable of at that configuration, regardless of AoA. 

[Ex-jazz]

Dear WMaker and Stoney,

If I understood this correctly, this calculation will give a whole plane Cl value, not a wing Cl value.

How this is relating directly to the wing area, I don't understand.

Thanks

[Stoney]

Dear WMaker and Stoney,

If I understood this correctly, this calculation will give a whole plane Cl value, not a wing Cl value.

How this is relating directly to the wing area, I don't understand.

Thanks

Because basically, the weight of the aircraft, divided by the wing area times dynamic pressure, gives you the Cl at that condition.  It alludes to the fact that the wing provides almost all of the lift for the plane, and that the fuselage et al doesn't contribute that much more.  

[Ex-jazz]

... but... but (tossing a book pages around)

In below of the best cruising speed case, the wing is producing more lift, than a plane weight due tail -lift.
In above of the best cruising speed case, the wing is producing less lift, than a plane weight due tail +lift.

 

[gripen]

The wing area is just a commonly accepted reference area for dimensionless lift coefficient despite the other parts of the airframe might be involved in the creation of the lift. 

[Ex-jazz]

The wing area is just a commonly accepted reference area for dimensionless lift coefficient despite the other parts of the airframe might be involved in the creation of the lift. 

Ah, I see. Thank you for the point out.

[boomerlu]

Stoney, I think I figured out how this works way back in our discussion about IAS vs TAS in P47 glide numbers.

Physics wise, the IAS speed will dictate how much lift the wing can generate. TAS is irrelevant for lift considerations if you have IAS available.

What TAS DOES affect though is your turn radius. You will pull a MUCH wider turn radius at high altitude for the same amount of Gs (i.e., lift).

[Angus]

In these stall tests, are there all figures calculated without power at the alt? Sorry if it is a silly question.

[FLS]

This would be the power on 1G stall. As you climb higher your IAS decreases and when it reaches stall speed you're at your altitude limit.