[Hummingbird]

Yo Yo it honestly looks like you're repeating exactly what I said, i.e. small change in Cl but big change in Cdi.

There is a change in Cl though, and I think 0.12 sounds reasonable considering the 50% increase in AR. IIRC there is a nice NACA graph that shows the difference I can find to illustrate it, I'll try locating it tonight.

[Hummingbird]

Another illustration:


Biggest part of why increasing the AR actually increases the overall Cl of the entire wing (not the airfoil polar) is the decreased influence of downwash over the wing: The higher the AR the smaller percentage wise an area of the wing is suffering a loss of lift due to downwash over the wing. Wing taper is another method used to also reduce this effect, and the Ta152 features both, and in combination I can definitely see it increase the CLmax by 0.12.

Note: I'm still looking for the side by side NACA illustration I was talking about, but I'll find it.

In addition to this an increase in Clmax of just ~.20 due to flaps seems extremely implausible (taking your 1.38 CLmax as benchmark), thus I really don't believe the FW AG figures to be with flaps.

As can be seen here the CLmax is increased by a value of between 0.8-0.95 in the area covered by flaps of the split type, so on an aircraft with flaps covering 50+% of the wing span I really can't see the overall Clmax rise by just 0.2.

Spoiler:

Spoiler:

On a seperate issue NACA also measured the CLmax of the F4U, F6F & P-51B at R= 5.8-6.1*10^4:

Clean:
F6F = 1.60
F4U = 1.45
P-51 = 1.39

Flaps down (60 deg):
F6F = 2.50
F4U = 2.21
P-51 = 1.92

Compared with the F6F or Fw190 the F4U obviously sacrificed abit of lift for increased stability with its bent wing shape, but the F6F's Clmax however straight up matches the FW AG figures of 1.58, the increased thickness of the wing raising it by 0.02.

Source:
https://engineering.purdue.edu/~aero...report-824.pdf

[Echo38]

Originally Posted by Hummingbird

the F4U obviously sacrificed abit of lift for increased stability with its bent wing shape

I always assumed this myself. Well, sort of. The wing was bent for the landing gear & prop clearance, not for stability. So, I always assumed that the designers of the Corsair sacrificed lift for shorter (and thus more sturdy) landing gear, by bending the wing. It seems obvious that the bent wing will have less lift than an unbent one, all else equal, because the lift vectors are not all parallel (and thus should have a weaker combined effect along the normal lift vector).

However, I recently read somewhere that, surprisingly, the bend in the wing somehow actually increased lift. Unfortunately, I don't remember where I read this, so this should (of course) be taken as simple hearsay, unless someone can help me out by providing a source.

[Hummingbird]

Originally Posted by Echo38

I always assumed this myself. Well, sort of. The wing was bent for the landing gear & prop clearance, not for stability. So, I always assumed that the designers of the Corsair sacrificed lift for shorter (and thus more sturdy) landing gear, by bending the wing. It seems obvious that the bent wing will have less lift than an unbent one, all else equal, because the lift vectors are not all parallel (and thus should have a weaker combined effect along the normal lift vector).

However, I recently read somewhere that, surprisingly, the bend in the wing somehow actually increased lift. Unfortunately, I don't remember where I read this, so this should (of course) be taken as simple hearsay, unless someone can help me out by providing a source.

Well as shown NACA measured the F4U's actual overall CLmax and it ended up lower than that of the F6F with its straight wing, and this loss in lift was attributed partly to the bend in the wing and partly to the addition of leading edge radiator intakes.

As for why the F4U featured the gull wing shape, I wasn't trying to claim it was for stability reasons, that was simply an added benefit. The real reason AFAIK was indeed to create the necessary ground clearance for the prop.

[Echo38]

I'm inclined to agree, and to assume that whoever originally wrote that the bent wing had better lift was mistaken.

[Curly]

Originally Posted by Hummingbird

In addition to this an increase in Clmax of just ~.20 due to flaps seems extremely implausible (taking your 1.38 CLmax as benchmark), thus I really don't believe the FW AG figures to be with flaps.

On a seperate issue NACA also measured the CLmax of the F4U, F6F & P-51B at R= 5.8-6.1*10^4:

Clean:
F6F = 1.60
F4U = 1.45
P-51 = 1.39

Flaps down (60 deg):
F6F = 2.50
F4U = 2.21
P-51 = 1.92

Are you referring to CL max of the 2d airfoil sectionals in that report? As I have seen much lower CL max reported for full scale wind tunnel and flight tests of those aircraft. The F6F for example is shown to have CL max of 1.3-1.4 in the clean configuration and close to 1.7 with the flaps fully deployed.

I find it plausible that in take off configuration, the CL max of the Dora is 1.58. The full flap of the F6F nets you a gain of .4 or .3 Cl max. It seems well within the realm of possibility that in the half flap configuration of the Dora you’re getting a increase of Cl max by .2

NACA reports 829 and 1044 both provide interesting result for wind tunnel and flight testing CL max numbers.

Also are the illustrations you’re looking for in regards to aspect ratio and Cl located here http://history.nasa.gov/SP-367/chapt4.htm
Figures 56 and 57?

[Yo-Yo]

Originally Posted by Curly

Are you referring to CL max of the 2d airfoil sectionals in that report? As I have seen much lower CL max reported for full scale wind tunnel and flight tests of those aircraft. The F6F for example is shown to have CL max of 1.3-1.4 in the clean configuration and close to 1.7 with the flaps fully deployed.

I find it plausible that in take off configuration, the CL max of the Dora is 1.58. The full flap of the F6F nets you a gain of .4 or .3 Cl max. It seems well within the realm of possibility that in the half flap configuration of the Dora you’re getting a increase of Cl max by .2

NACA reports 829 and 1044 both provide interesting result for wind tunnel and flight testing CL max numbers.

Also are the illustrations you’re looking for in regards to aspect ratio and Cl located here http://history.nasa.gov/SP-367/chapt4.htm
Figures 56 and 57?

The first question about flaps down lift is: WHY FW USED EXACTLY 1.58 AS A CL FOR LANDING?
Any "lift margins" is not valuable for taildraggers because 3 point landing requires two simultaneous events: mild stall slightly beyond the AoA that corresponds three point attitide.
Admiiting that clean CL max is 1.58 and keeping in mind the diagram for landing speed calculation, one must admit that ALL ADDITION from the flaps is absolutely useless regarding the landing speed. Moreover, this situation tells us that touchdown will be not at the three point attitude but at 6-8 deg attitude, that is full nonsense.

[Hummingbird]

Originally Posted by Curly

Are you referring to CL max of the 2d airfoil sectionals in that report? As I have seen much lower CL max reported for full scale wind tunnel and flight tests of those aircraft. The F6F for example is shown to have CL max of 1.3-1.4 in the clean configuration and close to 1.7 with the flaps fully deployed.

I find it plausible that in take off configuration, the CL max of the Dora is 1.58. The full flap of the F6F nets you a gain of .4 or .3 Cl max. It seems well within the realm of possibility that in the half flap configuration of the Dora you’re getting a increase of Cl max by .2

NACA reports 829 and 1044 both provide interesting result for wind tunnel and flight testing CL max numbers.

Also are the illustrations you’re looking for in regards to aspect ratio and Cl located here http://history.nasa.gov/SP-367/chapt4.htm
Figures 56 and 57?

No I'm refering to page 324 with the whole planform.

[Curly]

Originally Posted by Hummingbird

No I'm refering to page 324 with the whole planform.

That would be a model with a perfect finish. Page 296-297 discusses the discrepancy between the results obtained with a model and full scale tests. It cites a .2 lower Cl in the real world vs a model. Owing to roughness, leakage, intakes and armament installations. Taking .2 off the Cl max of the model subsequently puts the model results in agreement with the wind tunnel and flight testing of the full scale f6f. Which provides a good analog for the 190's airfoil.

Page 20 of NACA 829 goes into effect of the finish and seal on a service wing. Based on what's presented there, I think it's doubtful that there is anyway that a service condition 190 A or D has a clean configuration Cl max of 1.58. Not to mention Yo-Yo has presented plenty of both direct and anecdotal evidence that supports his claims of Cl max ~1.38.

[Pilum]

In the IL-2 Sturmovik forum a member (II/JG17_SchwarzeDreizehn) was kind enough to make the following figure and text available from a full scale test of a Fw-190 in the Chalais Meudon wind tunnel outside Paris performed by Focke-Wulf in 1943.

As can be seen, the figure shows a Clmax figure of 1.3 at Re=4.6 M (See attached figure for "Leerlauf"=idling engine). It is reasonable to look at the idling curve since this represents the same conditions (comparing apples with apples!) at which a Clmax of 1.36 for Spitfire and 1.4 for Me-109 was measured by the British RAE.

I think it is good to reference the Spitfire and Me-109 to get things in context since these Clmax values are AFAIK more generally accepted while the Fw-190 Clmax has been the subject of some controversy both here and in other forums and figures as low as 1.17 and as high as 1.58 have been mentioned.

My attempt at translation of the attached text from Focke-Wulf Bericht 06006, page 12 conclusions (See attached excerpt):

“Conclusions

In the large Chalais-Meudon wind tunnel, a full sized Fw 190 was tested.

Without split flap deflection the Camax turned out to be 1.3 and with 58 degrees split flap deflection Camax=1.55. These values are 0.3 to 0.4 lower than those of smooth/polished models. The deviation is due to the influence of fuselage, supports, and deviations in the wing profile shape.”

So at the Chalais Meudon wind tunnel Re of 4.6M the Clmax is 1.3 which given that an IRL stall speed Re is around 6.4M makes a Clmax at that Re in the order of 1.35-1.4 quite plausible.

So, it looks like this report backs up Yo-Yo’s choise of a Clmax in the range of 1.35 to 1.38 quite nicely!