Moderators please lock this thread.

[Holtzauge]

That is because David Lednicer is an educated aeronautical engineer and he is correct to use the 2D data if he wants to determine what the aircraft will do.

And being a qualified engineer he also understands what is reasonable and what is unreasonable. To be honest Crump, your claim of a no flap Clmax of 1.58 for the Fw-190 is hurting our chances to get the Fw-190 FM fixed since it can simply be dismissed as unreasonable by the developers.

[LuftManu]

And being a qualified engineer he also understands what is reasonable and what is unreasonable. To be honest Crump, your claim of a no flap Clmax of 1.58 for the Fw-190 is hurting our chances to get the Fw-190 FM fixed since it can simply be dismissed as unreasonable by the developers.

Stop.

 

Just make this a summary of things that need fixing.

[Crump]

 

 

To be honest Crump, your claim of a no flap Clmax of 1.58 for the Fw-190 is hurting our chances to get the Fw-190 FM fixed since it can simply be dismissed as unreasonable by the developers.

 

It is not my claim.  It is a fact that Focke Wulf and Grumman both used as the clean wing CLmax of 1.58 for the same airfoil selection.  

It is most certainly not unreasonable, but rather what the engineer should expect to see in the design.

[Crump]

It is a fact that the engineers at Focke Wulf and the engineers at Grumman agreed to the clean wing Clmax of that airfoil selection.  It is also a fact one can run the math for the Reynolds number one should see at the stall speed of the Focke Wulf FW190 and see it gives good agreement with the measured results we have on the aircraft.

 

110mph IAS = 161fps = FW-190A3 stall speed as measured by the RAE.

 

RN = (161fps * 5.95ft) / .0001756 = 5.45 X 10^6 Re

 

That is just below the 6.0 x 10^6 Re found on the NACA airfoil data and right in the CLmax of 1.58 determined by both the engineers at Focke Wulf and Grumman for the same airfoil selection.

 

 

It works out exactly as the physics says it should work.  Why would that be unreasonable?

[Holtzauge]

It is not my claim.  It is a fact that Focke Wulf and Grumman both used as the clean wing CLmax of 1.58 for the same airfoil selection.  

It is most certainly not unreasonable, but rather what the engineer should expect to see in the design.

 

No, It is your claim which is based on you misunderstanding the table which JtD and Matt have already pointed out several times.

 

It is a fact that the engineers at Focke Wulf and the engineers at Grumman agreed to the clean wing Clmax of that airfoil selection.  It is also a fact one can run the math for the Reynolds number one should see at the stall speed of the Focke Wulf FW190 and see it gives good agreement with the measured results we have on the aircraft.

 

110mph IAS = 161fps = FW-190A3 stall speed as measured by the RAE.

 

RN = (161fps * 5.95ft) / .0001756 = 5.45 X 10^6 Re

 

That is just below the 6.0 x 10^6 Re found on the NACA airfoil data and right in the CLmax of 1.58 determined by both the engineers at Focke Wulf and Grumman for the same airfoil selection.

 

 

It works out exactly as the physics says it should work.  Why would that be unreasonable?

 

A reasonable Clmax is in the order of 1.3 to 1.4 for the Fw-190, not 1.58:  From the Chalais Meudon test of the Fw-190, Focke-Wulf Bericht 06006 page 12 (See attached excerpt):

 

My attempt at translation:

 

“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 a figure of 1.6 to 1.7 is possible, but only for models, just like your NACA reports excerpts show and the authors of the Chalais Meudon 06006 report conclude.

 

In case you missed it, the developers are currently basing their Clmax on a value of 1.17 as far as we know and the idea with this thread is to compile data for a joint community FM report to the developers. Anyway, since this will be a joint effort, you are of course free to keep on pushing for your value of 1.58 but AFAIK that figure has already been dismissed (and rightly so IMHO) by the developers so I hope we will find consensus around a lower more reasonable figure like 1.3 to 1.4 supported by the data we have so far.

[Crump]

 

 

“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.”

 

That is from the same report as the 1.17 impaired polars.... 

[Holtzauge]

That is from the same report as the 1.17 impaired polars.... 

 

Impaired in what way that affects the measurement of the Clmax?

 

Edit: Added clarification that the question concern Clmax not the polar.

Edited October 2, 2016 by Holtzauge

[Crump]

Impaired in what way that affects the measurement of the Clmax?   Edit: Added clarification that the question concern Clmax not the polar.

 

 

All of the polars in the report are taken at the same Reynolds number Holtzauge.  

 

I am the one who posted all of them....

[Holtzauge]

All of the polars in the report are taken at the same Reynolds number Holtzauge.  

 

I am the one who posted all of them....

OK, so you mean it's because of the Re that you think they are "impaired"? Well that is why I think we should suggest a Clmax higher than 1.3 to the developers since the Chalais Meudon results were measured at Re=4.6M while IRL 1G stall would be at around 6.4M. So I would not say they are "impaired": They just need to be interpreted seeing that the results stem from a test with a somewhat lower Re than wanted.

Edited October 2, 2016 by Holtzauge

[Crump]

 

 

These values are 0.3 to 0.4 lower than those of smooth/polished models.

 

That is true only if the design team did not do their job and made a mistake.  Read the NACA reports and do not just cherry pick a graph.  The full sized aircraft, airfoil data, and model all give good agreement.  That is a fundamental principle in aerodynamics.

 

If a design team sees that it is not giving good agreement, then something is wrong.  

 

The NACA reports detail some of things that a design team should look for to fix the problem and outline some common issues.

 

1.  Pressure Leakage from poorly fitted fairing - fixed by adjusting the fit of the fairing

 

2.  Surface roughness - The surface of the aircraft must be equal to or smoother than a piece of aluminum rubbed down with 380 grit sandpaper.  That is not a hard standard to achieve.  Go hit up a section of your car with 380 grit sandpaper and compare that to the painted unsanded finish.

 

3.  Loose fitting skins - You must be able to place a straight edge over the wing and have no wrinkling of the skin as the straight edge is rolled down the rib.

 

4.  Correct wind tunnel corrections and data - The corrections used to account for the inclination of the relative wind produced by the windtunnel jet and the interference caused by walls, supports, and measuring equipment must be correct. 

 

So, if the wind tunnel does not agree with the flight measurements.....it is a big red flag to the design team.  

 

Nicely illustrating that is the Grumman design team flight measurements confirming this principle.

 

 

 

 

 

 

 

Anyway, since this will be a joint effort, you are of course free to keep on pushing for your value of 1.58 but AFAIK that figure has already been dismissed (and rightly so IMHO) by the developers so I hope we will find consensus around a lower more reasonable figure like 1.3 to 1.4 supported by the data we have so far.

 

Well, Grumman got very good agreement between the airfoil data, model, and flight measurements for the same airfoil selection.  The assumption we are now making is that Focke Wulf was unable to do the same thing that every other engineering firm does when they design an aircraft which is achieve that same good agreement.

 

All of their performance data is predicated on that foundation of airfoil agreeing with airplane.  When we do our performance math using the wing area as a reference, we have to use the characteristics of that same wing otherwise, our F=Ma relationship is altered!

 

Additionally, if we want to assume Clmax of 1.58 represents stall speed in the landing configuration (60 degrees of Split flaps) we would have to go back to the drawing board and explain a few things.

 

1.  Why does the RAE have the Touchdown Speed equal to the 1G stall speed.  That means the Coefficients of Lift MUST be the same.

 

It is a fact that Focke Wulf used 1.58 as the coefficient of lift used to determine touchdown speed.  The lift formula will tell you that the Coefficient of Lift at the same speed must be equal.  Therefore if the 1G stall speed equals the Touchdown Speed, the Coefficients of Lift are the same and Focke Wulf used 1.58 as the clean wing Clmax just like Grumman did for the same airfoil selection.

 

It is a fact that Focke Wulf list's a Clmax of 1.58 on their aerodynamic data sheet for the design.

 

2.  Why does Grumman's clean wing CLmax of 1.58 agree with Focke Wulf's data for CLmax of 1.58?  One cannot be for a clean wing and the other for 60 degrees of split flaps deployed.

 

IF you are correct and Focke Wulf's only represents 60 degrees of split flap deflection then the only assumption one can make is that same airfoil selection requires 60 degrees of split flap deflection to achieve the same lift force once you cross the Atlantic.   Now we have to explain what happens to the benefits of split flaps because the airfoil/flap design no longer corresponds to any known aerodynamic data and no longer fits the mold for a split flap design.

 

If we just look at the information we have that Touchdown Speed in the landing configuration equals 1G stall speed, then everything follows normal design and performance calculation convention without having to invent explanations.  Once more, the data from the manufacturer agrees with all the Allied data.  It all crosschecks.

 

It is pretty much a wrap.

Here is the aircraft configuration for gliding and climbing condition.  It is all clean wing 1G stall Clmax, just power on and power off.

 

[Crump]

OK, so you mean it's because of the Re that you think they are "impaired"?

 

They do not represent any velocity the aircraft can fly at.....

 

 If we know from flight testing that it takes 1348Hp to achieve 305KEAS then let's run the math to see our velocity at 9hp.

 

Our parasitic component of drag has a direct relationship with velocity and changes at the cube of velocity.

 

Therefore we know:

 

Pr2/Pr1 = (V2/V1)^3

 

That formulation is right out the book for BGS system aircraft performance calculations using subsonic incompressible flow theory.  

 

9hp/1348hp = (V2/305)^3

 

9hp/1348hp = V2^3 / 305^3

 

305^3 * 9/1348 = V2^3

 

 189431 = V2^3 = 57 KEAS = 29 m/s

 

Wow....

 

Somebody found the rest of the report!!  Fantastic!

 

 

 

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.

 

 

https://forums.eagle.ru/showpost.php?p=2910954&postcount=2

 

Nice to know that Performance Math works!!!

 

8% agreement off a SWAG for specific numbers and 100% agreement in conclusion!!  Difference in specifics is due to my propeller efficiency assumption.

 

Given the surface roughness issues and Reynolds number, the polar gives excellent agreement with the 2D NACA data when compared under the same conditions.

Edited October 2, 2016 by Crump

[JG13_opcode]

If the Fw is porked, and if the Chalais-Meudon tunnel delivered bad/flawed results, or if those results are being misinterpreted by the devs, then in order to get it changed the report needs to spell it out in a logical process, leading to an inescapable conclusion such that there can be no other interpretation.

 

The back-and-forth that's happening in this thread is the same back-and-forth that didn't work in the past.

 

I intend to start from first principles and build the case up from there in whatever direction the data supports.

[Crump]

Look forward to your conclusions JG13_opcode.

Let me know if I can help.

[JG13_opcode]

Will do, thanks. 

[Crump]

Found something to calculate with

 

Clean Stall: 118 IAS = 190 km/h -> taken from http://www.wwiiaircraftperformance.o...-47-1658-D.pdf

Wing area : 196,98 ft^2 -> taken from internet

Weight (max. take off): 9667,27 lbs = 4385kg -> taken from this game

 

Vso = sqrt(9667 * 2/(0.002377*1.38*197) = 173ft/sec = 190 km/h = 118 IAS

 

The CLmax is here = 1,38

 

I am a little confused about the in-game spec about the weight? Is the 4385kg the max weight with both MGFF with 90 rounds and fuel tank with 524 litres, because other source say something about maximum 3,978 kg (8,770 lbs) weight? I think the 3,978 kg (8,770 lbs) mean the Focke without outer MGFF cannons or maybe not?

 

This source say a Fw-190 A-3 with all wing cannons has 3978 kg (8770 lbs) https://books.google.de/books?id=wm_YDAAAQBAJ&pg=PA36&lpg=PA36&dq=fw+190+a-3+8770+lb&source=bl&ots=wsBP5RcMQ6&sig=ONTeU84SGyuiEKeGuQmDWTYRJPE&hl=de&sa=X&ved=0ahUKEwj23rfB5bbPAhXIXiwKHZG3D80Q6AEIVTAJ#v=onepage&q=fw%20190%20a-3%208770%20lb&f=false

 

Vso = sqrt(8770 * 2/(0.002377*1.25*197) = 173ft/sec = 190 km/h = 118 IAS

 

The CLmax is here = 1,25

 

Other source say about 4100 kg (9039 lbs) same weight that a Fw 190 A-6 has with all wing cannons http://www.wwiiaircraftperformance.org/fw190a/190a5-performancetable.jpg

 

Vso = sqrt(9039 * 2/(0.002377*1.29*197) = 173ft/sec = 190 km/h = 118 IAS

 

The CLmax is here = 1,29

 

My probability calculation of the CLmax value is between 1,25 ~ 1,38. However I think we can forget the 1,38 value because I think the in-game weight is maybe wrong 4385kg (9667 lbs) with all wing cannons. The Fw-190 A-6 weight with 4x MG151 wing cannons is 4100 kg (9039 lbs) so a Fw-190 A-3 with 2x MG151 + 2x MGFF wing cannons will weight a little less or between 3978 kg (8770 lbs) and 4100 kg (9039 lbs).

 

Short said we are talking about a CLmax from 1,25 to 1,29 for a clean stall at 118 IAS.

 

Stall occurs at the same CLmax and not the same speed.  Speed will vary with weight.

 

There is no Position Error curve and the value for your 118mph IAS is straight IAS.  WNr 313 PEC curve shows about -4mph at 110mph IAS and EB-104 shows -6mph at 118mph IAS.

 

The Focke Wulf data of 1.58 gives us a stall speed of 104 mph EAS  Add back in a 1.5mph CEC and a 5mph PEC and you are right at the 110mph IAS speed the RAE recorded.  It is not a stretch.

[Crump]

OK, so you mean it's because of the Re that you think they are "impaired"? Well that is why I think we should suggest a Clmax higher than 1.3 to the developers since the Chalais Meudon results were measured at Re=4.6M while IRL 1G stall would be at around 6.4M. So I would not say they are "impaired": They just need to be interpreted seeing that the results stem from a test with a somewhat lower Re than wanted.

 

6.1 x 10^6 Re is a CLmax of ~1.7 on the 2D data.....

 

I do not think it should be a negotiation based on emotion.  It is hard science and Focke Wulf gives us their value at sea level on a standard day.  That value agrees with Grumman assesment.

 

This is a fairly common airfoil selection for World War II designers.  It was attractive because of its higher than average CLmax and low buffet zone.  The trade off was a harsh accelerated stall.  

 

Not giving the wing design its advantages means your game is modeling something else based on arbitrary emotion and not the facts.

Edited October 3, 2016 by Crump

[Holtzauge]

The back-and-forth that's happening in this thread is the same back-and-forth that didn't work in the past.

OK. You are the editor and I'm sure Crump will supply you with ample material to work with. I'll just refrain from commenting it to avoid the back-and-forth.

 

 

Look forward to your conclusions JG13_opcode.

 

Let me know if I can help.

 

 

Will do, thanks. 

 

Good luck with the report then!

[JG13_opcode]

I detect sarcasm in your well-wishing.

[MiloMorai]

https://forums.eagle.ru/showthread.php?t=174561&page=2

 

This thread should be read.

[JtD]

If the Fw is porked, and if the Chalais-Meudon tunnel delivered bad/flawed results, or if those results are being misinterpreted by the devs, then in order to get it changed the report needs to spell it out in a logical process, leading to an inescapable conclusion such that there can be no other interpretation.

 

The back-and-forth that's happening in this thread is the same back-and-forth that didn't work in the past.

 

I intend to start from first principles and build the case up from there in whatever direction the data supports.

The main points have been brought to the attention of the developers, they now have the full CM report thanks to SchwarzeDreizehn. Basically, all is set for a meaningful discussion. However, given that it's an exchange of PM's and not a face to face meeting, I'm sure it will take some time to reach a conclusion.

[Crump]

 

 

The first question about flaps down lift is: WHY FW USED EXACTLY 1.58 AS A CL FOR LANDING?

 

 

It is not the CL for landing....It is the CL used to determine Landing speed.  

 

 

 

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.

 

And that is simply not true.  Typically taildraggers are stalled in the 3 point attitude.  The effect of flaps is to increase camber and LOWER the stall angle of attack.

What is nonsense is believing that 1.58 is the stall point.  Simply put,,,,you would not land....you would crash. 

 

Using the clean wing CLmax makes perfect sense in a military aircraft with only a 10% buffer above the stall on approach.  You arrive at the runway with a safety margin and when the pilot flares, the lowered stall angle will help to set her down.

[Crump]

 

 

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. 

 

 

Only if you do not convert the speeds listed in the reports....

 

If you do not you will find it very difficult to align those speeds listed with POH indicated stall speeds.  If you do convert the speeds, then it all lines up!!  Your CLmax is 1.4 to 1.6 for the F6F which agrees with the NACA flight testing of the type.

 

http://www.wwiiaircraftperformance.org/f6f/f6f.html

 

Stall testing is not performed at sea level.....

[CUJO_1970]

The current FW CLmax is what? 1.17?

 

Should be at least 1.38? Do I have that correct?

[=EXPEND=13SchwarzeHand]

Yes, the currently implemented parameter is 1.17.

Edited October 4, 2016 by II/JG17_SchwarzeDreizehn

[Crump]

 

et'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.

 

 

What he has determined is that will be flow separation.  That does not mean the wing has reached CLmax or is stalled.  Most likely, it is the onset of the buffet zone.

 

That is why the math he is using is behavior math not aircraft performance math.  It does not determine what the edges of the envelope but only tells us how the aircraft will act IN the envelope.

 

You can ONLY determine CLmax by sticking an airfoil section into a wind tunnel and measuring the results.  

 

That exactly what the NACA did in compiling their airfoil data.  The NACA 2D data for the 23015 airfoil was determined thru empirical means....it was measured under known conditions and reasonable predictions can be made as a result. 

 

There is some NACA airfoil data that was determined thru mathematical analysis and published just after World War II.  The NACA 66 series comes to mind.  The result is a whole bunch of airplanes with problematic airfoils that do not act as predicted.

 

[Livai]

Here some Wind tunnel test about the Fw-190 wing. This is a .pdf file that download automatic by clicking the link below

BTW if you searching inside the pdf file about the Fw-190 use these shortcuts (FW wing and Fw-190) to find the information much faster

 

https://www.google.de/url?sa=t&rct=j&q=&esrc=s&source=web&cd=24&ved=0ahUKEwjemMmHnMHPAhVUF8AKHbnlBWo4ChAWCG8wDQ&url=https%3A%2F%2Fwww3.nd.edu%2F~ame40462%2FRiegelsAerofoilSections.pdf&usg=AFQjCNHRGDPflOLC5qeJePvIMQ10hOpq3g&cad=rja

 

[Edit] I see that this source contain data about the Wing Tip Airfoil (NACA 23009) that the Fw-190 had, too

Edited October 4, 2016 by Superghostboy

[Crump]

 

 

Here some Wind tunnel test about the Fw-190 wing. This is a .pdf file that download automatic by clicking the link below BTW if you searching inside the pdf file about the Fw-190 use these shortcuts (FW wing and Fw-190) to find the information much faster

 

Great read and good for people to understand some of the complexity.

 

 

 

 

Velocity = (2600000*.000156927)/5.95 = 67.4 fps or 46 mph.

 

Well, the FW-190 cannot fly at 46 mph so that is hardly the Clmax of the wing under conditions an FW-190 can fly.

 

Knowing the surface roughness and Reynolds number and using the measured 2D data, we can get an idea of where the CLmax should be.... 

[Crump]

That's a different profile.

 

 

Yep...It is the 66-series...the ones based on calculations alone I spoke about earlier.  They only measure drag of the FW190 wing under the premise of testing surface roughness due to poorly prepared aircraft surfaces.  The Focke Wulf did surprisingly well.

 

Here is the NACA 23015, which the Focke Wulf's root profile.

 

 

6.0 X 10^6 is the vicinity of stall speed for the FW-190.  Here you can see the CLmax is listed as 1.7.

 

The little "r" is standard roughness.  Standard roughness has a grain size of .28 mm and covers 5% to 10% of the wing are on the suction and pressure sides.  That is 320 grit sandpaper on the wing along the top and bottom of the leading edge and approximates moderate rime icing.

 

That data exactly matches the NACA standard profile for the airfoil:

 

 

Compare that .28mm grain height with the .007mm grain height of the paint on the FW-190 wing.

 

Under the same conditions, the FW-190 wing was less effected than any of the other wings tested.  That is one reason the design is not significantly effected by service conditions.  Kurt Tank designed it to operate from unimproved fields under combat conditions.

 

http://www.wwiiaircraftperformance.org/fw190/Fw_190_A-2_216_220.pdf

 

1.) With factory No. 220 no gain in speed was recorded from the high-quality, highly polished condition

 

2.) With factory No. 216, the gain in speed amounts to ~ 3 mph (5 km/h) at and above full throttle height.

 

This result shows that the normal painting should be regarded as adequate in terms of surface quality and that a further effort for surface improvement is hardly worth it.

 

 

 

The diagram included in Superghostboys paper is part of study of surface roughness in which several wing designs where painted without primer with paint to gauge average roughness height.  It does not show anything about CLmax of the wing.  It does give good agreement with the German results on testing surface preparation.

 

 

What it does show is that the emergency finishes the Germans used in the last months of the war did impose a very small penalty but not much at all.  They sprayed their aircraft with very minimal surface preparation and no primer simply applying a light coat of paint over the aircraft.  The saving in time vs performance loss was worth it.

Edited October 4, 2016 by Crump

[Crump]

 

 

These are the theoretical clmax values, from a simulation of a 190 D IIRC. No historical document or actual measurement though. Although this definitely does give an imdication of the direction the values should be adjusted to.

 

Those are Computational Fluid Dynamics calculations and the best thing you get outside of actual measurement...

 

Here is model VASERO used to calculate that data:

 

 

 

David Lednicer sent me that.  It is included in the article but you cannot really see it so well due the resolution and small size.

[Crump]

 

 

To be honest Crump, your claim of a no flap Clmax of 1.58 for the Fw-190 is hurting our chances to get the Fw-190 FM fixed since it can simply be dismissed as unreasonable by the developers.

 

Bottom line is it can only be dismissed if we:

 

1.  Throw out all the Valid Measured Data and Computational Fluid Dynamics Analysis and go with a calculated Clmax.  Of course that is exactly the opposite of all aerodynamic convention.

 

2.  We ignore what the designers say the CLmax of their design achieves.  The same CLmax used by another firm for the same airfoil selection.

[Livai]

Look at this unbelievable to read this! Every write about a Wing Area of 196,98 sq. ft  but they here have a captured Fw-190 A-3 with a Wing area of 177 sq. ft!!!! Overall very interesting data about climb and a lot more!!!!

 

[JtD]

177sqft wing area plus 20sqft are 197sqft like everywhere else!!!!Totally wrong climb data!!!!!Physically complete nonsense!!!!!

 

!!!!!!!

[Bert_Foster]

Its a civil magazine article published at the height of WWII ... I am sure the Relevant RAF testing reports were at the time secret..... Hardly surprising "Flight's" report was not technically accurate.

[Phenazepam]

Hello all!
I'm FM engineer from 1CGS and I worked with FW-190 flight model.

Does anybody have "P. Gross "Die Entwicklung der Tragwerkkonstruction Fw 190", Bericht 176, der Lilienthal Gesellschaft, 2 Teil, January 1944"?

Also if you have any additional references don't hesitate to send us. They will be helpful for future work with FW190A5.

[1PL-Husar-1Esk]

Light in the tunnel ?

[Livai]

Secret RAF testing reports, lol. Publish at the height of WWII that a captured Focke climb at 1220m with 15,5m/s and at 5334m with 16,6m/s is more than stupid. You can really read how suprised RAF was at the end of the article from the flying legend.
 
Has someone this book " The Focke Wulf FW190 by Gordon Swanborough and William Green "? There should be something about British AFDU report on captured Fw 190 A 3 Werk-Nr 5313.  Telling how much superior in the most aspects to a spit 5 and again the same Focke climb data like above!

Here another book https://books.google.de/books?id=NFTEPiyEiSsC&pg=PA107&lpg=PA107&dq=Focke+Wulf+FW190++British+AFDU+report&source=bl&ots=RNWrrmHNMZ&sig=CL6KFe-17ieYRywq9EmRUIuYuO8&hl=de&sa=X&ved=0ahUKEwiE17uq4cXPAhUD6xQKHWLcCWUQ6AEIJjAA#v=onepage&q=Focke%20Wulf%20FW190%20%20British%20AFDU%20report&f=false

Here again talking a similar language like this book above.

 

In the end what tell this us the same to similar and more data can be found inside some books later published that repeat and confirm the data from the article published at the height of WWII 1943.

Edited October 6, 2016 by Superghostboy

[Phenazepam]

I repeat: German sources are in first row for german planes.

[=EXPEND=13SchwarzeHand]

Does anybody have "P. Gross "Die Entwicklung der Tragwerkkonstruction Fw 190", Bericht 176, der Lilienthal Gesellschaft, 2 Teil, January 1944"?

 

Would you mind telling us what sort of information you are looking for in that specific report? Maybe there are alternative sources.

Edited October 6, 2016 by II/JG17_SchwarzeDreizehn

[L3Pl4K]

 

 

Hello all! I'm FM engineer from 1CGS and I worked with FW-190 flight model. Does anybody have "P. Gross "Die Entwicklung der Tragwerkkonstruction Fw 190", Bericht 176, der Lilienthal Gesellschaft, 2 Teil, January 1944"? Also if you have any additional references don't hesitate to send us. They will be helpful for future work with FW190A5.

Maybe you can tell the people, which documents do you have.

[150GCT_Veltro]

Light in the tunnel ?

 

We hope......we really hope.