Maybe the 190s arent wrong....or how to be really unpopular in 1 step

[gripen]

Originally posted by Angus
Two different wings means two different airfoils right?
So, CL should be different, or could at least.
Well, in reality. Not sure of that theory.

It does not matter if the airfoil is different. Exactly the same Cl will be required for a given lift regardless geometry or airfoil.

Try to think it this way; same amount of lift is needed at given wing area despite what ever is the geometry or airfoil, so the required Cl must be exactly same despite what ever is AoA or induced drag.

Basicly claiming something else is a physical impossibility because the lift and area remain the same.

gripen

[Crumpp]

The required Cl will be exactly same at given loading despite what ever is the geometry of the wing. There is no need to know the angle of attack.

Who is talking about Angle of Attack??

The graph shows various angles of attack but the importance is the differences in the CL due to wing efficiency and Aspect ratio.

Are you not smart enough to compare them at 0 angle of attack listed on the chart?

Do you need me to point that out?

Ok, Gripen, look on the chart at the line marked "0" and compare the influence of induced drag and Aspect Ratio.

All the best,

Crumpp

[Crumpp]

Gripen says:

The formula gives directly the exact and accurate required 3D lift coefficient (Cl not Clo) for any given loading at any given flight condition.

Yeah ok.

That is why NASA says:


 

Keep in mind that this equation for lift is only a very basic expression used to analyze the overall aerodynamic behavior of an entire vehicle. Engineers must use much more complex methods to do detailed design and analysis work.

http://www.aerospaceweb.org/question/aerodynamics/q0078.shtml

Throw some more nonsense in the thread, Gripen.

All the best,

Crumpp

[gripen]

Originally posted by Crumpp
Who is talking about Angle of Attack??

The graph you posted above shows Cl vs angle of attack, but we don't need to know it.

Originally posted by Crumpp

The graph shows various angles of attack but the importance is the differences in the CL due to wing efficiency and Aspect ratio.

We are not interested what kind of angle of attack is required for the Cl required. Required Cl and lift will be the same regardless what ever AoA is needed at given flight condition.

Originally posted by Crumpp

Are you not smart enough to compare them at 0 angle of attack listed on the chart?

There is no need to know the AoA to calculate required Cl.

gripen

[gripen]

Originally posted by Crumpp

That is why NASA says:

"Keep in mind that this equation for lift is only a very basic expression used to analyze the overall aerodynamic behavior of an entire vehicle. Engineers must use much more complex methods to do detailed design and analysis work."

It does not change the fact that the exact and accurate required Cl can be calculated easily with simple equation.

gripen

[Crumpp]

It does not change the fact that the exact and accurate required Cl can be calculated easily with simple equation.

Yeah ok.  Tell NASA.

It can right next to your theory on calculating wing efficiency off the wing twist distribution chart in Lednicers article.

All the best,

Crumpp

[gripen]

Originally posted by Crumpp
Yeah ok.  Tell NASA.

It can right next to your theory on calculating wing efficiency off the wing twist distribution chart in Lednicers article.

They allready know it as they know it in the Aerospace web. They also allready know that there is a strong correlation between e factor and lift distribution.

gripen

[justin_g]

Find the CL for an elliptical shaped wing with no twist and an AR of 5:
Now find the CL for a rectangular shaped wing with no twist and an Aspect ratio of 7:
What is the CL going to be using the standard forumla??

Of course it is the same, the formula is being used to calculate the CL required from the wing to create the required lift under a given set of conditions.

But what will differ, is the AoA, drag etc. of different wings. This is where the more advanced theories/calculations would be applied.



Say you need CL = 1 to meet a given set of conditions. From the chart, you can see that the C172 only needs about 7º AoA, while the BAC Lightning would need about 20º AoA to get CL = 1. The Lightning would have higher drag as a result of the higher AoA and lower AR, requiring more thrust etc...

[Charge]

"It is interesting Charge that more than one flight test notes this same characteristic."

That may be, but my point was that I wouldn't put much weight on anything that that document says without knowing something about the aircraft/s in question ie. Power/weight figures, wing profiles and planforms+area, flap configurations (if used in test) etc.

As I said: Merely implications of their true performance.

-C+

[Angus]

ehh, Gripen:
"It does not matter if the airfoil is different. Exactly the same Cl will be required for a given lift regardless geometry or airfoil."

I recall HiTech explaining something similar. Bloody confusing.

Lift is a constant, - if you look at it that way.

[F4UDOA]

Of course it is the same, the formula is being used to calculate the CL required from the wing to create the required lift under a given set of conditions.

But what will differ, is the AoA, drag etc. of different wings. This is where the more advanced theories/calculations would be applied.

Justin sounds like he knows much more than I do however I will piggy back onto his point.

The equation to detemine Clmax is so simple that it encompasses all aspects of wing design without having to know almost anything about the aircraft. It could be a cardboard box you are measuring.

The calculation I use is

weight * 391(air density at sea level) / 1G stall speed ^2 * Wing area

There is no mention of aspect ratio, airfoil type or wing tip shape, why? Because these things are already factored into the equation in the form of stall speed. If the wing creates enough lift and the bird is not to heavy it will not stall. That is the same for a delta wing, Spitfire or 190. As long as you know the wing area, weight and 1G stall (power off) you know how this aircraft will handle.

What it does not tell you is Induced drag from lift. However an aircraft with lower wing loading will not have to pull as hard as one with high wing loading so it will naturally create less drag.

[hitech]

f4udoa: Whats the 391 term?

HiTech

[moot]

Air density at sea level in hectograms/cubic yard?

[Crumpp]

As I said: Merely implications of their true performance.

True and I completely agree.  They are not useful for absolute performance.
Of course niether are performance graphs from different countries in the 1940's without a more detailed information than speed/climb.

The set up sheet for this particular report is available listing.

All the best,

Crumpp

[F4UDOA]

HT,

I'm an idiot!! 361 not 391!!

The 361 is air density multipier for sea level. At least this is how it was expalined to me.

If you remember Wells from the boards a few years back, I used to correspond with him to try to understand this stuff a little better.

For example

F4U-1

12,000 * 361 / 96mph^2 * 314
4332000 / 2893824
Clmax = 1.49

That is actually a little high but the stall is really closer to 96.5TAS.

From there you can multiply the square of the G factor * the 1G stall number to get the accelerated 3G stall.

3G Squared = 1.73 * 96MPH = 166.27MPH stall at 3G

When you compare these numbers to the actual numbers they are not just close they are amost exact. (the chart is IAS Knots)