Moderators please lock this thread.

[unreasonable]

 

 

 

 

 

Just to understand what you are saying this graph is saying: the scale on the left top graph is the same as the scale on the right top graph - ie an interval of 0.5, values 0.5-1.0-1.5? Which is the obvious reading.  The top left looks a bit like 1.6 but that is just poor handwriting or reproduction. 

 

Or are you saying the graph says 1.6? Not sure from your posts.

[Dakpilot]

Ok, where do we submit the report to the devs?

 

I kind of thought that the whole point of this thread was in the title....

 

Cheers Dakpilot

[JtD]

i am sure it can and it is a real shame, too.  

 

As far as the question of what is the CLmax for the FW190 clean wing, that is a done deal.

 

It is a fact it is 1.58.

You haven't provided a single source that shows that. You've misinterpreted figures, you have manipulated figures and you've made up figures. Just like you're turning an assumption of 1.5 for a structural strength calculation into an aerodynamically viable 1.6 in this case. If you think repeating your fake arguments over and over somehow builds a case, you're wrong.

 

Whether JtD realizes it or not is another question.

Your condescending stupidities are really annoying. A 1g stall is structurally completely different from a max g stall, and just because you're incapable of reading a whole sentence transcribed from a standard, doesn't mean someone else has difficulties understanding.

Fall A conditions once more for you, put clearly instead of standard transcription: maxg, min speed at which maxg are obtainable, clmax obtained from that. A complex load case defined by three parameters. It's not my fault that this information arrives scrambled at your brain purely as "clmax".

As opposed to what you keep blabbering about 1g clmax and high g clmax don't have to be the same since a) wing geometry changes under high g (as in fact discussed in this very report) b) high speed aerodynamics have different properties from low speed aerodynamics, first in terms of Reynolds numbers and second in terms of Mach related effects. The standard leaves room for that, even if in case of the Fw190, 1g stall and high g stall were calculated with the same clmax for simplicities sake.

 

Relevant at this point is that according to figure 3 of the presentation ~0.5° of wing twist are obtained at clmax, max g flight and ~400kph, ~0.7° of wing twist are obtained at ~750kph and max g, proving that the high speed stall of the Fw190 was calculated with wing root stalling first. Try to get that into your head and leave your stupid remarks out of this topic. If you want to insult me, send me a PM with the moderators in CC.

Edited October 14, 2016 by JtD

[CisTer-dB-]

I suggest you guys meet on TS and discuss with each other. It seam to me that this is a typing monologue and no one understand each other. Please for the benefit of all of us, bring a mediator if needed but meet up and listen to each other.   I am convince that something good can come out of it

 

o7

Edited October 14, 2016 by ATAG_dB

[Crump]

I suggest you guys meet on TS and discuss with each other. It seam to me that this is a typing monologue and no one understand each other. Please for the benefit of all of us, bring a mediator if needed but meet up and listen to each other.   I am convince that something good can come out of it

 

o7

 

Honestly, It would be kind of pointless.  This is pretty cut and dry if you understand aeronautical sciences.  

[CisTer-dB-]

Honestly, It would be kind of pointless.  This is pretty cut and dry if you understand aeronautical sciences.  

I do, but still I find it difficult to follow you Crump

[Crump]

I do, but still I find it difficult to follow you Crump

 

 

 

Any questions feel free to ask.  Just PM me.  I am sure you understand the special relationship that CLmax has to structural load limits.

 

Maybe this will help with understanding wing sections and how those load span distribution charts in Die Entwicklung der Tragwerkkonstruktion der Fw 190 Bericht Nr. 176 der Lilienthal Gesellschaft - 2 Teil vom Januar 1944 are read.

 

Span loading is nothing more than comparing lift coefficient to wing span.  A basic span loading chart is simply wing section lift coefficient / Wing Lift Coefficient over the  wingspan from root to tip.

 

 

Let's compare VASERO to Focke Wulf Measurements:

 

 

At the bottom corner is a span loading distribution chart calculated by VASERO.  VASERO is a Computation Fluid Dynamics analysis software.  

 

 

Same Span loading distribution from Focke Wulf measurements on the clean FW-190 wing:

 

 

 

Here is how these charts work and what happens to the sections.  The root airfoil determines our CLmax.  When it stalls and is not longer able to support lift load, that load is shifted outboard and that outboard sections coefficient of lift will then rise until it reaches the CLmax for the airfoil.  When that section reaches a CLmax of 1.58, it stalls and shifts the load outboard until the aerodynamic twist prevents the section from reaching CLmax.

 

 

 

 

You can see the FW-190 exhibits typical spanwise load distribution is pretty normal and falls right where it should between a rectangular wing and some taper!

 

[unreasonable]

Just an observation on these CLmax calculations based on a measured power off level stall speed - they are extremely sensitive to this number, since this term is squared and is the biggest number on it's side anyway. 

 

For example, for a W of 3,900kg (37,278N) assuming sea level:

 

167 kph = 1.58

178 kph = 1.40

195 kph = 1.17

 

 

ie a 7% difference in this term would account for all the difference between the Crump maximum case of 1.58 and the suggested 1.40. 

 

Given that stalling the plane is a process (as the stall moves along the wing) and that test flights at the time did not - I believe - have particularly sophisticated instrumentation, I wonder how much sense it makes to believe that precision is possible?

[Irgendjemand]

Just a reminder guys. DONT GET THE THREAD LOCKED:)

[=362nd_FS=Hiromachi]

Exactly, how about we just wait until some update comes from Phenazepam ? 

[JtD]

JtD, what is your interpretation of those charts? What is the clmax number you get out of those, if it's not the 1.58 that crump is saying.

What you can take away from the report is spanwise lift distribution and wing deformation for some calculated load cases. They contain no information regarding maximum lift coefficient, because the 1.5 used is input data.

 

Given that stalling the plane is a process (as the stall moves along the wing) and that test flights at the time did not - I believe - have particularly sophisticated instrumentation, I wonder how much sense it makes to believe that precision is possible?

It certainly was possible to fairly exactly determine stalling point, angle and lift coefficient. While not a stationary thing like in a wind tunnel, it's still not highly dynamic and the job was typically carried out by highly trained and experienced staff. So one can assume that they were aware of issues, would consider them and if possible, would correct or at least properly interpret them.

However, if you look at RAE Spitfire measurements and NACA measurements with the same type, maximum lift coefficients were determined to be 1.36 and 1.15. The British claim the low US figures are a result of wrong speed measurements. So on one hand this shows the whole thing isn't fool proof, but on the other hand it shows that differences of 0.2 were considered a mistake, and not uncertainty of measurement.

Edited October 14, 2016 by JtD

[WWChunk]

Exactly, how about we just wait until some update comes from Phenazepam ? 

 

I agree.  I'm sure the discussion between JtD and Crump can be carried on via PM, and not in this thread, which has turned into a small pissing match.  I appreciate the work both of you are trying to do, but the measuring contest needs to end.  You're both smart, congratulations. 

[LLv24_Zami]

Just a reminder guys. DONT GET

THE THREAD LOCKED:)

This.

[JtD]

It surprises me that anyone thinks I'm discussing things with Crump here. I'm not. Whatever I'm telling, I'm telling YOU and not Crump. But the more feedback I read, the more I understand no one even reads the posts. Good to know, saves me time.

Edited October 14, 2016 by JtD

[JG13_opcode]

For [Edited] sake. All I wanted was one thread, argument-free.

Moderators please lock this thread.

Edited October 14, 2016 by Bearcat

[Crump]

 

 

What you can take away from the report is spanwise lift distribution and wing deformation for some calculated load cases. They contain no information regarding maximum lift coefficient, because the 1.5 used is input data.

 

Tell me what you think is beneficial in examining span loading at Coefficients of Lift beyond CLmax?

 

There is none, JtD.  An aircraft is unloaded the moment it stalls. 

 

This is exactly what the report says VERBATIM:

 

 

 

Im A-fall würde das flügelende des vergleichsflügels gegenüber rumpfmitte von 3 auf etwa .2 hochdrehen.  Daraus ergibt sich eine ca belastung, die offensichtlich zum beginn der strömungsablösung im querruderberich führen dürfte Der A-Fall wurde mit einem Auftriebsbeiwert des flügels von 1.5 gerechnet, da hierfür bereits vergleichsuntersuchungen der FW190 vorhanden wraen.  Der übliche ca wert von `1.6 könnte bei der angenommenen verwindung gerade etwa erreicht werden infolge des elastischen aufdrillens Flügels und ohne erhebliche gefahr des abkippens;  der unverwundene flügel dagegen gerät durch das aufdrillen am flügelende in das überzogene anstellwinkelgebiet und kann daher keinen nennenswerten ca-gewinn mehr aufweisen.  Er ist aber dabei in erhöhtem maße der gefahr des abkippens ausgesetzt.

 

The FW-190 wing is measured data from other studies and not some arbitrary value pulled out of thin air.  Again, using a value above Clmax in a structural examination of an aircraft is simply a silly idea.  The physics dictate that the structure is unloaded at CLmax.

 

 

 

The A-case (for the comparison wing) was calculated with a wing lift coefficient of 1.5, since comparative studies of the FW190 already exist. The usual ca value of `1.6 (for the FW190 wing) could be achieved by the elastic twisting of the wing and without a significant danger of the tip stalling;  the untwisting of the wing by the twisting at the end of the wing into the critical angle of the angle of attack and therefore making any Ca profit insignificant (tip stalling explained for the reader).   However, it (the comparison wing) is moreover exposed to a risk of tip stalling.

 

 

You can align all of that with the reports span loading graphs.

 

The FW190 wing has a Clmax value of 1.6 and NOT 1.5:

 

 

Once more you can easily tell that a completely different wing being examined.  The aerodynamic twist is noted for both the FW190 wing and the Comparison wing design.  The FW190 wing fits the description of the wing designs unusual aerodynamic twist of 2 degrees until 80% span.  The comparison wing has an aerodynamic twist of 3 degrees to 50% span.  Completely different wing designs.....

 

 

Explain the coefficient of lift value of 1.6 (1.58) which matches both Grumman data and Focke Wulf's data for their own designs.

 

 

 

 

How do you explain:

 

1.  Examining a coefficient of lift on any wing that exceeds CLmax?

 

2.  The Fact the graphs are clearly labeled FW-190 wing and Comparison wing

 

3. The FW190 wing shows the span load Clmax of 1.6(1.58).

 

4.  The Comparison Wing show the span load CLmax of 1.5

 

5.  The aerodynamic twist depicted on the FW190 matches the wing designs aerodynamic twist

 

6.  The comparison wing has a completely different Angle of Incidence.  The only way to change Angle of Incidence is to remove the wing and reattach it.  Do you think that is what being depicted....the inboard section is wrenched off and somehow the outboard section gets 30% of its aerodynamic twist removed?

 

Can you answer these questions please.

For fuck's sake. All I wanted was one thread, argument-free.

Moderators please lock this thread.

 

 

Give him a chance to answer the questions, please!  He is being civil.

[SvAF/F16_Goblin]

Stop this! You two need to take this to PM.

[Crump]

 

 

 I appreciate the work both of you are trying to do, but the measuring contest needs to end.

 

 

I completely agree that all the filler and white noise is a total distraction from the issue.  Give him a chance though to answer publicly the legitimate questions placed before him.  It is important.

 

I have only told the community how a span loading chart is read and how to interpet the data.  You can look in aeronautical textbook and it will tell you the same thing.

 

 

 

It is important to know why JtD believes that Focke Wulf is examining span loading at coefficients of lift that exceed the design CLmax?

 

 

 

In Flight Mechanics, the coefficient of lift will ALWAYS be a value below CLmax once you come off the stall line.

 

Edited October 14, 2016 by Crump

[Crump]

The FW-190A  design maximum coefficient of lift is 1.58.

 

At 8580 lbs weight for WNr 313 at measured at a stall speed of 110mph IAS let's see what our Equivilent Airspeed has to be in order to achieve a Clmax of 1.58 and does that Equivalent Airspeed fall without our range of possible speeds.  WNr. 313 is a Type II fighter variant with 4 wing cannon.  The weight difference is due to the fact the RAE used natural Petroleum AVGAS which weighs ~7.48lbs per imperial gallon.  C-3 fuel is closer to jet fuel and is a light oil with a distinctive burnt coal smell and weights ~7.86 lbs per imperial gallon.

 

 

 

 

 

Using the BGS system and a conversion factor for knots...

 

EAS means sigma (density ratio) = 1

 

Velocity in KEAS = SQRT{295 (8580lbs) / (1.58 * sigma * 197 ft^2)

 

Velocity equals = 90.2 Knots Equivilent Airspeed

 

90.2 KEAS * 1.15 = 104 mph EAS.

 

110 mph IAS - 104 mph EAS = Correction error of 6 mph.  

 

104 mph EAS + .5 mph CEC = 104.5 mph CAS = 104.5 mph CAS + 5.5 mph PEC = 110 mph IAS as recorded by the RAE on WNr 313.

 

The only absolute relationship of stall vs angle of attack easily returns a stall speed that is well within the margin of error found in airspeed measurement.

 

Given the fact the Allied data and German data both give good agreement regarding stall data we can gain some insight referencing both sets of data.

 

The 1G level power off stall point for a Type II fighter is the same airspeed as the Landing speed.  Both equal 110 mph IAS.

 

 

A Clmax of 1.58 represents the clean configuration (gear up, flaps up) design CLmax.  This can be used to find the correct stall speed for any weight or configuration of the aircraft.

 

Edit'd to add the allied PEC curves to show all data is in agreement.

Edited October 14, 2016 by Crump

[Crump]

Here is a Focke Wulf Position Error Correction which shows us the Position Error Correction as measured by the RAE shows an accurate trend.

 

 

 

 

[Crump]

If you cannot follow all of that...

 

 

 

The only absolute relationship of stall vs angle of attack easily returns a stall speed that is well within the margin of error found in airspeed measurement.

 

 

That is the conclusion.  The only absolute relationship that never changes is stall vs angle of attack.  Focke Wulf's determination of a Clmax for the clean wing, power off stall vs angle of attack relationship gives good agreement with measured results.

[unreasonable]

Even assuming that speed is correct, how do you know that the weight was 8,580lbs when the plane landed?

 

Was there ammunition on board? Or equivalent weight? The report does not say - or does it?

If the numbers are run assuming no ammo on board, (as I would think was likely), the CLmax is 1.50, assuming your speed.

 

But wait - what about the fuel weight at the end of a test flight? Again we do not know, but surely this would be a fraction of the full weight - let us assume the tanks are half full. Without ammo and 50% fuel the plane would be about 10% lighter.  Recalculate the CLmax, assuming your speed - 1.42!

 

There are just too many unknowns in the calculation for it to be particularly convincing. 

[Crump]

 

 

Just an observation on these CLmax calculations based on a measured power off level stall speed - they are extremely sensitive to this number, since this term is squared and is the biggest number on it's side anyway.    For example, for a W of 3,900kg (37,278N) assuming sea level:   167 kph = 1.58 178 kph = 1.40 195 kph = 1.17     ie a 7% difference in this term would account for all the difference between the Crump maximum case of 1.58 and the suggested 1.40. 

 

You have NO Position Error Correction or Compressibility correct in those speeds. 

 

167 kph = 1.58

 

You have to add back in the Position Error and Compressibility Error

 

167kph + 10 kph PEC + 1 CEC = 178kph or ~180kph Stall speed using a CLmax of 1.58

Even assuming that speed is correct, how do you know that the weight was 8,580lbs when the plane landed?

 

Was there ammunition on board? Or equivalent weight? The report does not say - or does it?

If the numbers are run assuming no ammo on board, (as I would think was likely), the CLmax is 1.50, assuming your speed.

 

But wait - what about the fuel weight at the end of a test flight? Again we do not know, but surely this would be a fraction of the full weight - let us assume the tanks are half full. Without ammo and 50% fuel the plane would be about 10% lighter.  Recalculate the CLmax, assuming your speed - 1.42!

 

There are just too many unknowns in the calculation for it to be particularly convincing. 

 

 

It is corrected mathematically so these philosophical arguments are not valid.

[JG13_opcode]

Give him a chance to answer the questions, please!  He is being civil.

Indeed, he is. JtD is not the person who is frustrating me in this thread.

 

Here is a list of quotes:

Yeah, that is nice piece of dancing around you printed there JtD.  I hope you enjoy "looking smart" and feeding your ego.

 

You got it JtD or do you need me to walk you through it?

Whether JtD realizes it or not is another question.

Would you call these quotes "being civil"?  Because I would not.

Coming from Crump, lol!

 

And you. I've never seen you make a substantive or positive contribution to a thread, ever. Your entire reason for being on this boards appears to be to just follow Crump around like a lost puppy and clutter up threads.

 

Can you honestly say that this post I've quoted had any purpose other than antagonism?

[JG13_opcode]

Edited the OP.

[Crump]

 

Indeed, he is. JtD is not the person who is frustrating me in this thread.

 

Here is a list of quotes:

 

1.  "Yeah, that is nice piece of dancing around you printed there JtD.  I hope you enjoy "looking smart" and feeding your ego."  - Guilty - I honestly could not reach any other conclusion in reading his post.  It is obvious that he did not know the Flight Mechanics relationship of CLmax to any structural investigation of a wing design or how the details that allow one to interpret a span loading chart.  It was intended to emphasize the fact we do not need more white noise to muddy the waters and he should check his conclusions.  My reply laid out the enormous hurdles his interpretation had to overcome.   There is somewhat of language issue in communicating with the Devs that makes eliminating barriers to communication imperative.

 

I do not claim to know everything but at least construct an argument that fits the norm for aircraft and flight mechanics.

 

 

I apologize to JtD, you, and anyone else for making it.  To me, his entire post could be summed up as "Crumpp must be wrong and I will prove it at all cost" to the point of fabrication and I will take responsibility for reacting to it.

  

2.  "You got it JtD or do you need me to walk you through it?"  Not Guilty.  Honest question and one I more than willing to help him out.  JtD is a smart guy and has much to offer.  He is a Safety Engineer and not formally trained or educated in Flight Mechanics.  That does not invalidate his input but the constant arguments just to attempt to prove me do nothing but add constant static to the background of any discussion.  If I am wrong, I simply admit it.

 

3.  "Whether JtD realizes it or not is another question."  Not Guilty.  Honest observation.  Given JtD propensity to oppose everything simply based on me being in the conversation and goes back to my offer to help him out.

 

Would you call these quotes "being civil"?  Because I would not.   I can see how the first one fed into the other two comments.  All I can say do is take responsibility and pay better attention to what others perceive.  That being said, I am not an arrogant jerk and this is a BBS over the internet so we are missing some important clues in interpersonal communications.  Keep that in mind please and I am always open to PM's or questions.  If you ask, "Did you mean ....."  I will be happy to clear up any confusion.

 

Edited October 14, 2016 by Crump

[Crump]

Thank you for taking the time to present a summary of the article for the English speaking audience. I would, however, like to add/correct some points to the ones you've presented.

 

You are welcome.  Your input is actually valuable when you are not trying to win an argument just to win it.

 

What it basically says is that the new construction ("Schalenbauweise"/stressed shell) offers higher torsional stiffness and damage resistance at a reduced weight compared to a conventional construction ("Holmbauweise"/spar construction).

 

I got that from reading it but semi monocoque wing construction vs truss was not answering the question, "What is the Clmax of the FW-190 wing?"

 

Figure 2 shows the conventional, comparison wing. You description is right, but the article doesn't refer to the Fw190 wing.

 

I never said it refered to the FW190 wing...only that it defined the conditions of the wings being compared.  That is clean configuration without flaps and that the Coefficient of Lift was set to a value of 1 as representing flight.

 

Fall A is the maximum lift coefficient obtained at the lowest speed at which maximum permissible g can be obtained, Fall B is the lift coefficient at maximum g at 0.8 maximum permissible ram pressure (~0.9 max. permissible speed (vne)). Both cases are safety relevant structural strength calculations. 

 

That is the definition of CLmax. 

 

The values are theoretical and caTrmax is defined as "the maximum stationary lift coefficient of the aircraft normalized to the wing area".  Which is the Clmax of the wing design over the area of the wing design.  It is NOT a standard value of 1.6!  Normalized in aerodynamics usually means we are taking non steady state flight such as a stall or turning flight and turning it into a comparative value as IF it was steady state flight.  It is the "normalized" Value of Clmax in a turn to seperate the now divided components of lift.  It looks suspiciously just like the lift formula...

 

L/q*S = Normalized Lift

 

In this case, the lift coefficient for figure 2 was normalized using the math definition to a value of ONE from the measured Clmax value of 1.6(1.58).   That is why Figure 2 has the note CaFL = 1. 

 

 

 

multiply (a series, function, or item of data) by a factor that makes the norm or some associated quantity such as an integral equal to a desired value (usually 1).

 

https://www.google.com/webhp?sourceid=chrome-instant&ion=1&espv=2&ie=UTF-8#q=normalized+definition

 

Typically, a caTrmax of 1.6 was used at the time.

 

Never heard of this "standard value" and it does not fit how aircraft work.  That is an extremely high CLmax that MOST airfoils could never achieve.  There is absolutely no benefit to calculating anything structural above Clmax of the wing design.   It does however fit very nicely into the fact both Grumman and Focke Wulf report after extensive investigation and measured data collection similar value of  CLmax in their wing designs using the NACA 23015/NACA 23009 wing design and the NACA measured data on the airfoil. 

 

Explain where you came up with this "standard value" of 1.6, please?  The FW 190 wing design in the article is clearly labeled with a Clmax of 1.6(1.58) as found in Focke Wulf GmbH studies while the comparison wing is calculated at a Clmax of 1.5.  It looks very much like you did not know what the term "normalized" means and decided on your own that it must be a "standard" value to me.

 

 

All results are calculated, none are measured. Assumptions are on the conservative side in order to not jeopardize safety margins.

 

 The report most certainly does state that the values for CLmax for the FW-190 wing are measured.  

 

It states:  "Der A-Fall wurde mit einem Auftriebsbeiwert des flügels von 1.5 gerechnet, da hierfür bereits vergleichsuntersuchungen der FW190 vorhanden wraen.  Der übliche ca wert von `1.6 könnte " Which translate to english is "The A-fall was calculated with a lift coefficient of the wing of 1.5, since comparative studies of the FW190 already exist. The usual ca value of `1.6 could be achieved...... "

The key points in figures 3 are:

Less torsion in the Fw190 wing than there is in the conventional wing.  If you mean more torsional rigidity, I agree. Focke Wulf accepted a weight increase to preserve the agility of the design.

Safe stalling behaviour of the Fw190 wing when compared to the conventional wing, because even under high loads the Fw190 wing stalls root first, where the conventional wing starts to stall tip first (A-Fall). Agree.

Slight reduction in maximum obtainable lift coefficient for the conventional wing, since the large wing area near the root can no longer be brought up to clmax. Agree

 

About your summary - engineers don't do structural safety calculations for a clmax of 1.5 and then have the aircraft fly around with a clmax of 1.58. Doesn't happen.  Smart alec comment meant to antagonize me which is why you got one back.   First of all, a value of 1.5 is the comparision wing and NOT the FW-190 wing.  The FW190 wing's CLmax is 1.6(1.58).  Most importantly,  CLmax is the gold standard and all structural limitations will be either AT Clmax or below it.  This "standard value theory of 1.6" does not fit in any aircraft design convention and looks fabricated simply to argue with me. 

Edited October 14, 2016 by Crump

[II/JG17_HerrMurf]

Edited the OP.

We just can't have nice things (in the forums). The game, however, is great.

[JG13_opcode]

We just can't have nice things (in the forums). The game, however, is great.

 

The game is, in my estimation, not great but good.  It has serious flaws.

 

It can be great.

 

Make IL2 Great Again.

Edited October 14, 2016 by JG13_opcode

[CUJO_1970]

I appreciate all the data shared in these threads, by all parties.

 

There sure does seem to be a whole lot of information that supports a CLmax at or very close to 1.58 for the FW-190.

 

It seems we are flying the aircraft around at 1.17 right now. That's a big difference in handling and agility.