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


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Old 09-24-2016, 01:37 PM   #1
Kwiatek
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Default Question for Yo-Yo about Fw 190 Clmax and cAoA?

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)

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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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Old 09-26-2016, 02:14 PM   #2
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Originally Posted by Kwiatek View Post
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)

\




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

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Old 09-28-2016, 09:22 AM   #3
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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?



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

Last edited by Kwiatek; 09-28-2016 at 10:40 AM.
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Old 09-28-2016, 08:14 PM   #4
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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.
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Ніщо так сильно не ранить мозок, як осколки скла від розбитих рожевих окулярів
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Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

Last edited by Yo-Yo; 09-28-2016 at 08:21 PM.
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Old 09-29-2016, 02:04 PM   #5
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Ummmm, maybe this should be pinned, so YoYo doesn't have to repeat it every time someone finds this document?
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Old 09-29-2016, 03:52 PM   #6
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Quote:
Originally Posted by Solty View Post
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.
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Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me
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Old 09-29-2016, 07:19 PM   #7
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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.



The FW AG clmax figures seem completely legit.
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Old 09-29-2016, 07:43 PM   #8
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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.



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.
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Ніщо так сильно не ранить мозок, як осколки скла від розбитих рожевих окулярів
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Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me
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Old 09-29-2016, 08:14 PM   #9
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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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Old 09-29-2016, 08:45 PM   #10
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Originally Posted by Hummingbird View Post
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.
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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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