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Question for Yo-Yo about Fw 190 Clmax and cAoA?


Kwiatek

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If it is not top secret information i really appreciate Yo-Yo if you could answer some data about Fw 190 wing polar.

 

These is German document which show Fw 190 Clmax without aircreew effect clmax 1.2 and cAoA 15.5 deg ( clean configuration)

 

\

 

 

index.php?app=core&module=attach&section=attach&attach_rel_module=post&attach_id=23034

 

Some find that these test got error and Clmax and cAoA got should be higher for Fw 190 ?

 

I wonder DCS D-9 is based on which data and how looks Clmax and cAoA for these bird?

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If it is not top secret information i really appreciate Yo-Yo if you could answer some data about Fw 190 wing polar.

 

These is German document which show Fw 190 Clmax without aircreew effect clmax 1.2 and cAoA 15.5 deg ( clean configuration)

 

\

 

 

index.php?app=core&module=attach&section=attach&attach_rel_module=post&attach_id=23034

 

Some find that these test got error and Clmax and cAoA got should be higher for Fw 190 ?

 

I wonder DCS D-9 is based on which data and how looks Clmax and cAoA for these bird?

 

 

The problem of this document (amongst the community, at least) is that nobody presents the essential information from the full report, so a lot of speculations was born on forums regarding the conditions of these tests, etc.

 

First of all, the air speed during the main tests conducted for the lift, drag and, thus, polars was 36 m/s. It was directly specified in this report, so Re was 4.6*10^6.

This is lower than the lowest IAS in flight.

 

As it is known, there are two opposite tendencies for the CL_max vs Re and M, but in the low M area the tendency of increasing CL prevales. So, at the 1g stall IAS we can suggest (and the results of NACA tests with NACA 230 trapezoid planform wing (clean wing without a fuselage similar to FW 190 wing) prooves this suggestiona. Though it's impossible to directly compare these two sources due to different Re/M coupling, the whole plane and idealised polished model w/o fuselage, so the TN 1044 was used for the further estimations using F6F-3 as the most closest (but not full!) analogue.

 

Finally, 1.35-1.38 seems to be right for this plane regarding it's known flight performance, original wind tunnel data, similar wind tunnel and flight tests of NACA.

 

 

The great problem known for the forum battles around the CLmax is that some people mix in one pile men and horses, apples and oranges, trimmed lift of the whole plane, where the fuselage eats sufficient part of lift changing the circulation along the wing, and the high AoA lift is lowered with the stab negative force - and the isolated wing, the wing with the different planform, etc...


Edited by Yo-Yo

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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Thx Yo-Yo for your replay. I think i understand problem with estimated Clmax for Fw 190 expecially based on too low velocity tunnel test and lack of other tunnel test and data directly for Fw 190.

 

I wonder one thing casue in some Fw 190 data German put 1.58 Clmax value?

 

zn7x2d.jpg

 

I checked by simple stall speed test in game for D-9 and got she has about 17 deg cAoA ( at least no less)?


Edited by Kwiatek
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No, CL=1.58 can not be maximal clean CL. Never.

 

I won't use even any external sources and will try to unveil my old analyses of this table.

This famous table provides somewhere more questions than answers... and we need to apply some math to estimate what we have behind these Ca_XXX.

 

First of all, we do not know exactly what kind of results it contains - measurements, calculated data or mixed, so, we will crosscheck within the table, checking our suggestions.

 

1. The table contains polinomial (2nd order or parabolic) estimations for the two parts of the polar for different conditions - high speed flight or climb. If they are obtained from the tests, we can say that they are for the different Re, M and cooling settings (as we can see, cooling is listed as a part of drag). Both of them are plotted (for 190 A) using table specified coefficients on the chart. Obviously both polars are for the clean airplane.

 

2. Then, we tried to plot the couples of Ca/Cw from the table. Obviously, the points DO NOT belong to the clean polars - and they are very far from it, though they seem to be at the same curve but shifted right. It means that these points ARE NOT FOR THE CLEAN plane, whatever the indexes mean, but for the same configuration.

Ok, let's try to suggest that it is START configuration - flaps 12 grad and undercarriage down. The table contains drag areas for these additions - 0.09 and 0.55. As ususal, to reduce it to Cw they must be divided to the F area.

The yellow curve is low speed (climbing) polar shifted right to this value. The points seem to be in the right place now.

 

But we have not touch this 1.58 magic number, yet... We just showed that these Ca/Cw couples are neither for the full flaps nor for the clean configuration.

The right approach for it will not be in guessing "who is who" in these couples (they are removed from consideration) and generally it will be comparative amongst different plane types.

 

First of all, we have F_kl/F ratio that presents relative flaps area. Then we have additional drag area due to the flaps deflection. And, that is very interesting (!), we have CL max = 1.7 for the TA152 H. Sweeping away the thought that FW mods for the 230 NACA rose clean CL to this value, just suggest that the difference is due to the different flaps area, for example...

 

Then we recall, that A, D had the same wing, TA152 C had almost the same wing with slightly extended wingspan. The flaps of all these modifications had the same span. TA152 H had very different wing with increased flaps area and SPAN (that is important for the further considerations).

 

So, it's time for the estimation itself.

 

Let's take a look at the table at the bottom of the chart:

flaps area is calculated using F and F_kl/F ratio. Then k is a ratio of F_kl to F_kl for A model.

The first check is to compare drag area added by the flaps at the 60 deg deflection where they act more like an airbrake, and yes - the absolute area flaps ratio between 190A and 152H is about 1.42 and the drag area ratio is 1.4.

THen, for the further estimations the scale drawings of the wings were used.

The wings of 190A and 152H have almost the same flap chord ratio (average through the flaps span) and center position along ther wingspan (to make sure the circulation distribution is almost the same). So, in this case deltaCL additions fo the lift coefficient of the clear plane will be proprtional of the fracture of portions of the wing affected by flaps.

THe estimation gives about 1.25 for 152H. Then, deltaCl for 190A is 1.58-1.35(DCS estimation)= 0.23, then CL max for the 152 H (presuming the CL max of a clean plane is the same as for 190A) is 1.35 + 1.25*0.23 = 1.64.

 

THis is close to 1.7 specified in the table... I do not think it should be exactly 1.7 because the exact CL max for the clean TA 152H is unknown - different planform, slightly different airfoil.

392655247_FW190tablecheck.thumb.gif.90048bd87de60e3fde27a16dff4f4420.gif


Edited by Yo-Yo

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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Ummmm, maybe this should be pinned, so YoYo doesn't have to repeat it every time someone finds this document? :)

 

I guess, I never posted the above mentioned results... I think they could stop the Crystall Ball divination what these CaXX mean.

 

And, by the way, this table does prove that A series and D series using the same wing have the same polars excluding, for sure, non-induced drag due to the different fuselage.

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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An increase in aspect ratio will also increase the CLmax mainly due to minimising the effect downwash has on the wings overall lift, so the Ta-152H will have a higher CLmax for that reason. A value 0.12 higher due to an AR increase of 6 to 8.94 doesn't seem off at all.

 

downwash.gif

 

The FW AG clmax figures seem completely legit.

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An increase in aspect ratio will also increase the CLmax mainly due to minimising the effect downwash has on the wings overall lift, so the Ta-152H will have a higher CLmax for that reason. A value 0.12 higher due to an AR increase of 6 to 8.94 doesn't seem off at all.

 

downwash.gif

 

The FW AG clmax figures seem completely legit.

 

As you can see from your own drawing, the effect you are trying to claim right works significantly only for low AR. High AR wings do have different dCL/dAoA slope but the CLmax is not affected because the slope asymptotical is very close to the ideal slope.

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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c

As you can see from your own drawing, the effect you are trying to claim right works significantly only for low AR. High AR wings do have different dCL/dAoA slope but the CLmax is not affected because the slope asymptotical is very close to the ideal slope.

 

Well 0.12 is not really significant, however it is noticable just as the visible difference between an AR of 5 and 7 on NASA's illustration. The truly significant effect is the reduction in lift induced drag.

 

That having been said the whole idea behind increasing the wing span on the Ta152 was to substantially increase lift so that the aircraft could effectively maneuver at higher altitudes, the difference in AR between the Fw190 and Ta152H was afterall 2.94.

 

Anyway Wiki actually has a nice explanation on this:

"Roughly speaking, an airplane in flight can be imagined to affect a circular cylinder of air with a diameter equal to the wingspan. A large wingspan is working on a large cylinder of air, and a small wingspan is working on a small cylinder of air. For two aircraft of the same weight, employing different wingspans, the small cylinder of air must be pushed downward by a greater amount of force than the large cylinder in order to produce an equal upward force. The aft-leaning component of this change in velocity is proportional to the induced drag.

 

The interaction between undisturbed air outside the circular cylinder of air, and the downward-moving cylinder of air occurs at the wingtips and can be seen as wingtip vortices."

 

Same reason the F-14 featured swing wings, to increase the lift when needed in part by increasing its wing AR :)

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c

 

Well 0.12 is not really significant, however it is noticable just as the visible difference between an AR of 5 and 7 on NASA's illustration. The truly significant effect is the reduction in lift induced drag.

 

That having been said the whole idea behind increasing the wing span on the Ta152 was to substantially increase lift so that the aircraft could effectively maneuver at higher altitudes, the difference in AR between the Fw190 and Ta152H was afterall 2.94.

 

Anyway Wiki actually has a nice explanation on this:

"Roughly speaking, an airplane in flight can be imagined to affect a circular cylinder of air with a diameter equal to the wingspan. A large wingspan is working on a large cylinder of air, and a small wingspan is working on a small cylinder of air. For two aircraft of the same weight, employing different wingspans, the small cylinder of air must be pushed downward by a greater amount of force than the large cylinder in order to produce an equal upward force. The aft-leaning component of this change in velocity is proportional to the induced drag.

 

The interaction between undisturbed air outside the circular cylinder of air, and the downward-moving cylinder of air occurs at the wingtips and can be seen as wingtip vortices."

 

Same reason the F-14 featured swing wings, to increase the lift when needed in part by increasing its wing AR :)

 

You are not right again. The main goal was to slightly decrease wing loading and significantly decrease induced drag. The main problem of medium AR (5.5-6) wings at the high altitude is an unwanted coupling of normal maneouvring IAS, high Mach number and low engine power because it works above FTA.

Decreasing of induced drag leads to significantly higher L/D ratio that is equivalent to engine power increasing. So, you gain more rate of climb and better sustained turn even without changing the engine.

Generally, you simply move the "coffin corner" to the higher altitude.

AR itself has very low effect either to CL before stall or to the CL cryt, especially beyond the AR 5...6, though it has progressive effect to the induced drag and L/D ratio.

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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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.

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Another illustration:

EffectsofAspectRatio.png

 

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.

 

gLYh5.jpg

 

 

NaLKgRd.gif

 

 

 

 

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/~aerodyn/AAE514/Spring%202011/naca-report-824.pdf


Edited by Hummingbird
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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.

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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.

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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?

[url=http://history.nasa.gov/SP-367/chapt4.htm][/url]

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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?

[url=http://history.nasa.gov/SP-367/chapt4.htm][/url]

 

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.

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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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?

[url=http://history.nasa.gov/SP-367/chapt4.htm][/url]

 

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

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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.

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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! :thumbup:

Fw190Bericht06006page12.png.595ba8fe4c902ca5f85085e30063a416.png

Fw190ClmaxFigure18cut.thumb.png.5404fcff9fcf0c93a444c07a34d274e0.png


Edited by Pilum

 

Old Crow ECM motto: Those who talk don't know and those who know don't talk........

 

http://www.crows.org/about/mission-a-history.html

 

Pilum aka Holtzauge

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Keep in mind that the same influences will occur with the P-51 which smooth/polished model recorded a Clmax of 1.39 at NACA, whiilst the F6F recorded one of 1.60.

 

Just for reference.

 

FW AG's figures are probably for a smooth & polished airframe also, which was std. procedure anyway AFAIK.

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  • 6 months later...

 

“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! :thumbup:

 

The FW-190 simply cannot fly at that velocity so how can anyone conclude that is the CLmax??

 

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.”

 

This is not a retraction of earlier data but a statement that all the data agrees!!

 

1.3 + .3 or .4 = 1.6 to 1.7

 

In other words, the 1.58 for the clean configuration FW-190 CLmax is correct. That is why Focke Wulf used it. This is part of the parametric studies the wing design article refers too, btw.

 

:smilewink:

Answers to most important questions ATC can ask that every pilot should memorize:

 

1. No, I do not have a pen. 2. Indicating 250

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What Clmax is actually being used in the DCS Dora ?

 

It affects the stall characteristics of the aircraft, and overall situations where maximum lift generation is required.

 

It's very important for us glider pilots, on pair with the Reynolds number.

Flight Simulation is the Virtual Materialization of a Dream...

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Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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