CLmax Calculator

October 30, 2016

Apart from that, I'm still wondering how I best would proceed to design a simple aircraft to given simple specs without giving myself away to chance excessively. Or is this the nature of things anyway?

You start with an idea of how to solve the problem, and progressively go complex&costly. You would, for instance, built and investigate models, mock-ups and components before finally going for the full scale fully functional aircraft.

October 30, 2016

You start with an idea of how to solve the problem, and progressively go complex&costly. You would, for instance, built and investigate models, mock-ups and components before finally going for the full scale fully functional aircraft.

All the while keeping an eye on the agreement of your original design specifications. The two most important for any aircraft are the design coefficient of lift and CLmax of the design. Those two parameters form the basis of the design performance you expect. They define all your landing, take off, and workload parameters. If you do not have your CLmax correct then your landing gear design, tail design, control surface design, in fact nothing for landing or takeoff performance will be correct. The stability and control of the aircraft will not be as designed either.

If they are not attained then your design will not perform the task as intended.

So, you keep an eye on the agreement. If you keep the conditions correct, you should see good agreement. If you do not and the conditions are the same, then you start looking at finish, leakage, construction, etc..until you find the problem and get that agreement.

In fact there is so much of aircraft design that depends on those two coefficient's of lift that being unable to predict them simply make us unable to design aircraft to anything specific. Our cropdusters vs fighter vs cross country day tripper vs heavy cargo transport from short unimproved airstrips.....would all be products of chance!!

October 31, 2016

You would, for instance, built and investigate models, mock-ups and components before finally going for the full scale fully functional aircraft.

Given the lift coefficients define the use of your plane, I am surprised to see that Focke Wulf GmbH (and others?) keep on listing the "theoretical" CLmax in their spec sheet when in fact that value should be something to be determined experimentally on the respective airframe.

It is also of notice that it seems neither Heinkel, Focke Wulf, Junkers or Messerschmitt had (if they felt like having one, they would have owned one) an own "full scale" wind tunnel. They either built the whole plane (is that efficient?) and found a volunteer to tell them how the crate is performing or, often enough long after the airplane is in service, they rented out facilities at Rechlin or used pirated installations in "their part of Europe" and got soem second opinions. If they did an awful lot of "modelling", I don't know what to make of that. I imagine a long wait list to put something in the wind tunnel at Rechlin. And there should be a paper trail documenting "true" lift values. Is there such a thing?

At least in the case of the Fw-190, I've learned that they originally had 2 different wings for that aircraft, differing about 1 meter wingspan and the shorter wing never really made it into service. Thus, there is some sort of try and error. But I'm not sure of the context there.

I mean, if one could really only measure lift parameters with the whole "real" aircraft, building planes like the He-177 must have been very interessting for sure. There was absolutely no wind tunnel in the world to put that one in. But given it was built to mornonic specs anyway and never got suitable engines, there was never much hope in it. Heinkel himself hated it.

October 31, 2016

Given the lift coefficients define the use of your plane, I am surprised to see that Focke Wulf GmbH (and others?) keep on listing the "theoretical" CLmax in their spec sheet when in fact that value should be something to be determined experimentally on the respective airframe.

Holtzauge has given an explanation in post 76 above - or are you talking about different sheets?

October 31, 2016

Question: I note that the CLmax for the Ta 152C and E is also listed as 1.58, even though the profile has changed from 23015.6 to 23016.8, yet Crump made a great song and dance about NACA 23016 being a completely different profile. How could the CLmax be identical for different profiles, when the root profile is supposed to be the main determinant of the wing's (therefore the aircraft's) CL max?

Edited October 31, 2016 by NZTyphoon

October 31, 2016

Holtzauge has given an explanation in post 76 above - or are you talking about different sheets?

the 1.5 to 1.6 2D Clmax figure from the Fw-190 report SchwarzeDreizehn was kind enough to provide comes from a structural analysis by Focke-Wulf so should not be confused with an aerodynamic analysis.

So, you say Focke Wulf put there the data from structural analysis? For the layman, when it is all about the aerodynamic analysis, what is the structural value good for to put there? Where is the added value for the reader of this sheet? Why not puting the aerodynamic value there that suposedly gets discovered iteration by iteration of modeling and testing? I mean, it would make the aerodynamic value/analysis awfully interessting. However no one seemed really to care about that "true" value. If I were them, I'd put the "true" value in gold over every following drawing that my future is based upon... But maybe it's just me.

How could the CLmax be identical for different profiles, when the root profile is supposed to be the main determinant of the wing's (therefore the aircraft's) CL max?

In the context of my question, this is strange indeed. Or do these different 230x profiles indeed have a similar "structural" (is there a proper name for that?) CLmax?

October 31, 2016

How could the CLmax be identical for different profiles, when the root profile is supposed to be the main determinant of the wing's (therefore the aircraft's) CL max?

Because it is the same airfoil...

Only the Ta-152H had a different airfoil selection.

Focke Wulf Ta 152 NACA 23015.3 NACA 23009

http://m-selig.ae.illinois.edu/ads/aircraft.html

It is not that Focke Wulf did not know what they were doing.....

October 31, 2016

At least in the case of the Fw-190, I've learned that they originally had 2 different wings for that aircraft, differing about 1 meter wingspan and the shorter wing never really made it into service. Thus, there is some sort of try and error. But I'm not sure of the context there.

It had a smaller wing and was faster. The handling was better with more wing area. The small wing actually did make it into combat and a handful of FW-190's to see service had it. It is like 5 or 10 aircraft, I could look it up or Milo will rush in here and tell us. The pilots liked the bigger wing so it won out.

Amazing that they could predict the performance changes in their design to get the desired performance.....huh?

'Holtzauge', on 18 Oct 2016 - 13:13, said: the 1.5 to 1.6 2D Clmax figure from the Fw-190 report SchwarzeDreizehn was kind enough to provide comes from a structural analysis by Focke-Wulf so should not be confused with an aerodynamic analysis.

There is no "structural vs aerodynamic" Clmax.

There is only the "aerodynamic" Clmax of the design.

Simply put.....The aerodynamic CLmax determines the structural limits. You cannot exceed it because the structure is unloaded the moment you reach the aerodynamic Clmax.

October 31, 2016

Because it is the same airfoil...

Only the Ta-152H had a different airfoil selection.

Focke Wulf Ta 152 NACA 23015.3 NACA 23009

http://m-selig.ae.illinois.edu/ads/aircraft.html

It is not that Focke Wulf did not know what they were doing.....

Wrong: the airfoil for the Ta 152C and E is clearly listed, in the document provided by Crump, as 23016.8, or doesn't Crump understand how to read his own material? Let's spell it out, shall we?

NACA 230 = an airfoil in the 230 series: 16.8 = NACA 23016.8:

15.6 = NACA 23015.6

CLmax for Fw 190 A-8, A-9, D-9 D-12 listed as 1.58,

yet CLmax for Ta 152C, Ta 152E with different airfoil =...1.58

So, how come Focke-Wulf got their airfoil profile data so wrong?

October 31, 2016

Wrong: the airfoil for the Ta 152C and E is clearly listed, in the document provided by Crump, as 23016.8, or doesn't Crump understand how to read his own material? Let's spell it out, shall we?

That is not the profile thickness....it is an area.

2D airfoil profiles do not have 3D areas....

I am betting it is the wetted area of the wing for form drag at zero lift as it is very close to the area of the wing.

October 31, 2016

I am betting it is the wetted area of the wing for form drag at zero lift as it is very close to the area of the wing.

It is the area of the wing without any control surfaces....

October 31, 2016

Given the lift coefficients define the use of your plane, I am surprised to see that Focke Wulf GmbH (and others?) keep on listing the "theoretical" CLmax in their spec sheet when in fact that value should be something to be determined experimentally on the respective airframe.

It sounds to me as if you overestimate the importance of lift coefficients. Sure the thing needs to fly, but then, how exactly do lift coefficients define the use of the plane? You can't use an aircraft as a bomber if it doesn't have a lift coefficient of 1.2345678? Didn't happen. Maybe it didn't fullfil some requirments - too long take off distance - optimize flaps. Too short range - fine tune aerodynamics. Too slow climb - get a bigger engine - and so on.

Focke Wulf gives the figures used for performance calculation on their aerodynamic data overview sheet. These values are assumptions with different degrees of validation through testing. If they calculate their landing speed using 1.58 and the take off speed with 1.4 - what else would they put into a table? Now they would at some point need to check if these values are OK, and they did that for instance with the testing at Charlais Meudon.

Full scale tunnels, depending on the size of the aircraft, existed for instance in Berlin (central research centre) and Dessau (Junkers), as already said. It's certainly possible to dig up material in particular from Adlershof.

The change of the wing of the Fw190 certainly isn't related to a misjudgement of clmax or anything, it's just that they felt that maybe it is a good idea to use a bigger wing, and when they tried it, it was. You can bet that they made some theoretical analysis before they built it, which already suggested it would be a good idea. However, the decision to actually use the 18.3m² wing was made only after the trials. Which is how things work(ed).

October 31, 2016

It had a smaller wing and was faster. The handling was better with more wing area. The small wing actually did make it into combat and a handful of FW-190's to see service had it. It is like 5 or 10 aircraft,

Considering the V5 and V6 tested a large and small wing and all subsequent Fw190As had the larger wing sounds like another reading comprehension problem like the removal/installation of the fuselage fuel tank depending on the mission to be flown someone was emphatic about.

The small wing was ~10kph slower than the big wing but more maneuverable.

October 31, 2016

Considering the V5 and V6 tested a large and small wing and all subsequent Fw190As had the larger wing sounds like another reading comprehension problem like the removal/installation of the fuselage fuel tank depending on the mission to be flown someone was emphatic about.

The small wing was ~10kph slower than the big wing but more maneuverable.

October 31, 2016

It sounds to me as if you overestimate the importance of lift coefficients.

Not really at all. In fact, being able to estimate them and once you have picked the airfoil for the design achieve good agreement is extremely important.

You can't really even design an airplane without being able to do that, LOL.

5.4.4. Airfoil Selection Criteria Selecting an airfoil is a part of the overall wing design. Selection of an airfoil for a wing begins with the clear statement of the flight requirements. For instance, a subsonic flight design requirements are very much different from a supersonic flight design objectives. On the other hand, flight in the transonic region requires a special airfoil that meets Mach divergence requirements. The designer must also consider other requirements such as airworthiness, structural, manufacturability, and cost requirements. In general, the following are the criteria to select an airfoil for a wing with a collection of design requirements: 1. The airfoil with the highest maximum lift coefficient ( max Cl ). 2. The airfoil with the proper ideal or design lift coefficient ( d Cl or i Cl ).

You need to indicate the locations of all parameters on the airfoil graphs. Solution: By referring to figure 5.22, the required values for all parameters are determined as follows: Cli Cdmin Cm (Cl/Cd)max o (deg) s (deg) Clmax Cl (1/rad) (t/c)max 0.2 0.0045 -0.03 118 -1.5 12 1.45 5.73 9%

The locations of all points of interest are illustrated in Figure 5.22.

http://faculty.dwc.edu/sadraey/Chapter%205.%20Wing%20Design.pdf

Wing design is ALL about being able to achieve the design coefficients of lift.

November 1, 2016

Knew you would bite on the speed and wing comment and ignore producing any data on the on the small wing seeing combat.

November 1, 2016

Question - if the 15.6 on the data table is a wing area of some sort, as the table headers suggest, perhaps it is the area which Fw used with 1.58 to calculate Vmin?

Comes out as plausible enough if you try it in the calculator, especially if you knock of a 100kg or so for missing ammo. Then recalculate the CLmax using the same Vmin but the wing 18.3 figure - 1.38

Just emphasizes another problem - unless you know exactly what wing area was being used in any calculations you still get the same proportional degree of uncertainty.

So is there any evidence that 1.58 was used in conjunction with 18.3 m^2 in Fw documents? Or are we just assuming this number 18.3 for some reason?

If Fw was using 15.6 as a wing area to calculate one variable, why would it not use it for all of them?

Not meant as rhetorical questions btw.

Edited November 1, 2016 by unreasonable

November 1, 2016

Knew you would bite on the speed and wing comment and ignore producing any data on the on the small wing seeing combat.

Please stop this tedious multi-thread ankle biting.

November 1, 2016

That is not the profile thickness....it is an area.

2D airfoil profiles do not have 3D areas....

I am betting it is the wetted area of the wing for form drag at zero lift as it is very close to the area of the wing.

Mr Speaker, I'm willing to admit that I got that wrong, so I withdraw and apologise. Dang, it proves I'm human after all!

November 1, 2016

Not meant as rhetorical questions btw.

Neither were mine, but thanks to all of you guys, I got really good answers that I'm sure of helped to get me to the bottom of this issue and I think I have the solution to most of our arguments here.

Let me explain my line of thinking and all of you are most welcome to comment/correct/[...] on it. This is gonna be a longer post (my apologies) so I will start out with the

Abstract for the nervous amongst us:

1) Crump, JtD, Holtzauge and most of the rest are actually on the same page, if it wasn't for some unnessesary words between them.

2) Unreasonable, in your spreadsheet you really made an error, the one I was making for years and still did in all those pages of this thread.

3) Surface imperfections are not required to explain the deviation from plane lift values compared to published CLmax of a profile.

4) Calculating the performance enveloppe of a plane is an exact science that still requires iterations to perfect it. Then you book time in a wind tunnel.

So, point 1, why are Crump and Holtz (and others, but for shortness sake I'll stick to those to "different opinions") on the same page?

First of all let's look what CLmax is:

It is the maximum lift value that a given profile can produce. It does so on an infinitely long wingspan at a given Reynolds.

But how does a real world airfoil look like? That one is a mix of (most often) two different profiles, a more lift generating one at the wing root at a higher angle of incidence and then a slightly less lift producing profile towards the wing tip, featuring a more or less (as in the case of the 190) reduced angle of incidence.

What happens now if I have such an installation on an aircraft and I'm approaching stall speed?

With increasing AoA, the wing will produce more lift, that one being cancelled out by slowing down. In any case I'm travelling up the CL curve on the published lift charts of the profile(s). This goes on until the onset of stall.

Now, what is happening at this precise moment?

The wing section that reached maximum AoA first will lose the lift. And it is very safe to say that it will do so at the published CLmax value.

What does that mean for the rest of the wing? As the loss of lift is occuring in two discrete section (just the starting points on the left and the right wing airfoil, it is a very very small part of the wing), the entire rest of the wing, owing to washout, is, especially at the wing tip not at the same CL (and certainly not CLmax). But you can calculate rather straight forward how much lift it will produce.

Thus:

"When published CLmax of the wing root section is reached, basically the entire wing is not there yet and is producing a CL less than the aforementioned Clmax".

Exapmple:

I have a plane with 12 m wingspan

Chord 1.5 meters

a fuselage 2 meters wide (i want to sell to fat people too)

Zach's001 profile for the whole wing, yielding a CL of 0.2 per degree AoA up to a CLmax of 1.6.

I have 2 degree (constant) washout.

The wing is rectangular, such as in the "Hershey Bar" Piper Cherokee 180.

I'm flying with that one, and I'm approaching stall. This will happen at 8 degree AoA. Why? I'm gaining .2 CL per degree, and this (almost, and in my profile most certainly) up to the value of 1.6.

Thus: 1.6 / .2 = 8

So, I'm travelling at a given speed (voluntary homework: calculate the speed of the aircraft at that point when, having a fat passenger next to me, the total weight is 900 kg, and the Reynolds Nr. being as you deem suitable) what is the lift I can expect?

If I have 2 degrees washout then that means that on average the whole wing is 1 degree behind the wing root section! And when the wing root section stalls at CLmax 1.6, and 8 degrees AoA, and the average wing is 1 degree AoA behind that, then the wing is at an AoA

8 - 1 = 7

So, my wing for practical purposes now stalls at 7 degrees, not at 8!!

If I look up my profile table and can see that I'm losing .2 CL per degree, lo and behold with a CLmax of 1.6, I'm stalling at a CL of 1.4!!!

This is my conclusion 1), that both "Crumpists" and "Hotzists" are in fact threading on the same thing. It's just that it took me all the way from Holtz' posting #76 until here to make some substance of it. Crump provided basic knowledge to prove Holtz right as well.

This in mind brings us to point 2), unreasonable making the same error on his preadsheet as I made over years.

The error is that we both assumed an uniform wing.

Looking at the example above, you can see that most of the wing is not stalling. The aircraft is approaching stalli This is more than semantics. As illustrated above, it is not just wing area (that had to be corrected in a second round) that matters, but it is also important at what configuration all sections of that wing are. In the case of the Fw190, IIRC the entire wing section in front of the ailerons are at an AoA being 2 degrees less that the root. The formula doesn't reflect that but is averiging out the wing CL produced when the root section reaches 1.58!

In the light of my wing above and looking at the charts, deducting 1 degree AoA gives you a pretty good guess about the lift produced at stall speed. AND it puts JtD's and Holtz's guess of ~1.4 as CL for lift into a very plausible perspective.

i feel calling that lift coefficient "CLmax" did us a big disservice, as it let us mix up CLmax of an airfoil, wing lift coefficient at stall speed what is actually happening aerodynamically on the wing at stall speed.

In this sense, you shouldn't call it "CLmax" on your spreadsheet, unreasonable. Using your spreadsheet, we calculate something else than max lift on a theoretical airfoil and therefore we should not call it the same. (We could call it "CL_fokked"?).

This is bringing us to point 3.

In light of the differences between CL max and maximum CL of your full airctaft, adding a lift penalty for "unattainable" lift values give you just what we have as a Fw190 now. A maximum CL of 1.17 or something sad like that. You maybe have that if a Yak did some writing on your wings with his cannons. But not with an airworthy design.

Point 4, all this being exact science.

If I am just remotely correrct with my resoning here, you can see that a designer still can know for sure (if he is worth his salary) what he is drawing up as a future aircraft.

The CLmax of a profile I'd say is maybe as important of a factor as compression ratio in an engine. If you want performance, you should be literate to handle it.

Then again, JtD's remark

You start with an idea of how to solve the problem, and progressively go complex&costly.

made me dig deeper.

Again, he's right of course. Or, more than right enough. Guessing... one can guess if your spouse still likes you, but guessing in a hundert years old science, nah, must be a diffrent sort of guessing.

I found a good calculator for assessing and building aircraft:

http://www.pca2000.com/en/software/ads-validation.php

I guess there are even better ones, but you can see the precision with which you can predict the performance (not: behaviour) of your plane. No rainbow colored 3D models (they are better to investigate the behavior) needed for that.

And there I also found a great description of the itearative nature of the design process, taken from this document:

http://www.pca2000.com/en/rapports/algorithms.pdf

First, what variables do we need for this process:

Then we start the first round of iterations:

then the second round:

After that, the process is no longer iterative, as we determined the respective values sufficiently.

When you had the guys in the shop build you came up with, then you can bring it to the wind tunnel.

While you guys may or may not agree with each other, all in all, you are guys telling the same story. And I'm sure the devs must have some sort of tool like the one I mentioned in the link that will tell you the "real" performance envelope by entering wing data. This way you get your "correct" maximum lift value for the whole plane. Moreover, you can even assess whether the guys operationg a wind tunnel made a good job or not by looking at their data.

Z

November 1, 2016

This in mind brings us to point 2), unreasonable making the same error on his preadsheet as I made over years.

The error is that we both assumed an uniform wing.

I assumed nothing of the kind! I assumed that:

1) The proposition from the aerostudents.com summary of aeronautical formulas was true:

"Also, it is interesting to notice that the minimal speed an airplane can have, can be calculated, if the maximum lift coefficient is known: W = L = CLmax / ( 1/ 2 * ρ * Vmin^2 * S) "

2) In which case, the Clmax can also be calculated if the (edit) other terms are known instead. Ie that the formula is an equation not a one way function. (Forgotten the technical term )

Nothing you have said in your post indicates that these are mistaken. So if you do think there is an error, please specify what it is - is the original proposition incorrect? Is my algebra incorrect? Is the spreadsheet wrong?

The formula cares nothing for the shape of the wing, only the total lift created and the area.

It seems to me that the only reason you think that my use of (edit) it is mistaken is because it can easily show a figure other than 1.58!

Going back to the 1.58 - the only agreement I can see is that this number might be applying to a wing CLmax since it is a plausible reduction from the 2D section CLmax for which we have documentation.

The disagreement is the difference between the wing and whole plane CLmax. Just as NACA discusses - or were the NACA engineers in the grip of error too? These variations would be an exact science if you knew all of the myriad effects of gun ports, wheel wells etc before hand: but they did not.

My other concern - as raised in my last post just before NZTyphoon's confession (+1 for him for taking it on the chin) is that even if we agree that 1.58 was used by FW in various calculations - perhaps including the Vmin one as above - do we know what wing area they used? I ask because I originally assumed this was not a problem but now I wonder since there are various ways to define the reference area used for the calculation.

Edited November 1, 2016 by unreasonable

November 1, 2016

First of all let's look what CLmax is:

It is the maximum lift value that a given profile can produce. It does so on an infinitely long wingspan at a given Reynolds.

Having discussed why I do not think I am in error, I want to go back and say where I think the cause of your misunderstanding is coming from - right there in that quote.

Indeed you can find other sources that state that CLmax is a property of airfoils - but then there are also lots that say it is a property of aircraft.

Here is one - who am I to argue with Stanford? (Well why not, but this is their opinion )

http://adg.stanford.edu/aa241/highlift/clmaxest.html

It is not just about the airfoils - it is about what all the other stuff does to the lift generation.

One more post to come .... all of this cutting and pasting is hard work. . cannot get rid of box below, please ignore.

Edited November 1, 2016 by unreasonable

November 1, 2016

Last one on Z's megapost: I agree that Crump, Holtzauge, JtD all agree that aircraft CLmax =/= 2D airfoil CLmax, but they do disagree about how much. In so far as I can understand Crump's argument at all, he is simply denying that anything like gun ports etc will have a measurable impact. Everyone else, including NACA, thinks that they do.

As to Crump providing basic knowledge - well here is what he had to say about it in post 38 in the "NACA indicates..." thread:

"That is the take away from examining the wing design, construction, and airfoil selection Focke Wulf used. Anything that does not fall within that range is simply incorrect. A complete measured polar of the aircraft will produce an airplane CLmax of ~.2 less than the 2D data.

Hint:

If you do the Re math for the NACA 23015 airfoil at speeds the FW-190 could fly.......

It comes out to 1.6. 1.6 - .2 = 1.58

Craziness, huh?"

Craziness indeed.

November 1, 2016

Mr Speaker, I'm willing to admit that I got that wrong, so I withdraw and apologise. Dang, it proves I'm human after all!

You going to put that up on your signature too or is that just a judgment reserved for yourself alone?

........@

"That is the take away from examining the wing design, construction, and airfoil selection Focke Wulf used. Anything that does not fall within that range is simply incorrect. A complete measured polar of the aircraft will produce an airplane CLmax of ~.2 less than the 2D data. Hint: If you do the Re math for the NACA 23015 airfoil at speeds the FW-190 could fly....... It comes out to 1.6. 1.6 - .2 = 1.58 Craziness, huh?"

Completely out of context but OK buddy, obviously it is something that fits your agenda...

Question - if the 15.6 on the data table is a wing area of some sort, as the table headers suggest, perhaps it is the area which Fw used with 1.58 to calculate Vmin?

Comes out as plausible enough if you try it in the calculator, especially if you knock of a 100kg or so for missing ammo. Then recalculate the CLmax using the same Vmin but the wing 18.3 figure - 1.38

Just emphasizes another problem - unless you know exactly what wing area was being used in any calculations you still get the same proportional degree of uncertainty.

So is there any evidence that 1.58 was used in conjunction with 18.3 m^2 in Fw documents? Or are we just assuming this number 18.3 for some reason?

If Fw was using 15.6 as a wing area to calculate one variable, why would it not use it for all of them?

Not meant as rhetorical questions btw.

Great theory.....really....

Completely wrong and nonsensical but it does sound good.

The reason for having the area of the wing minus the area of the control surfaces is to determine the sticked fixed and stick free stability and control characteristics of the design.

But go with what you came up with!!

Edited November 1, 2016 by Crump

November 1, 2016

The formula cares nothing for the shape of the wing, only the total lift created and the area.

Exactly. The "error" is certainly not in the math. The "error" is that one creates a good source of mixing up the published CLmax for a profile with the maximum lift coefficient an airfoil can have. Those "CLmax'ens" have not much to do with each other. The "theoretical" one says when the profile achives maximum lift. The other tells you the starting point where a real world airfoil, consisting of various, different profile sections, has achived maximum lift. The numbers therefore cannot be equal. Them not being equal is also not dependent on a suposedly defective finish.

And yes, just one report, no matter from where it is, is nowhere near enough substance to overthrow current practises in aero engineering. Nature (welcome to the club, biologist ) has published some papers that "were not that good" as well.

Taken together, I stand by my original assessment that your spreadsheet is very useful nevertheless. One just has to be aware of the limits or what it tells you and what it doesn't. that's all.

And besides, the link I gave you shows an example of a software doing for routine what we are trying to reproduce here. That program is an elaborate and more detailed version of your spread sheet and it is just one of several software tools to do such.

None of them have a parameter "for the practical deviations of real world profile sections compared to published values". It is just not a topic at all. Nowhere. Therefore, I still must assume that published values are (like the ominous 1.58) used. So the NACA report was not considered important enough (even for NASA) to "include it" in current design principles. So, why should I?

November 1, 2016

So is there any evidence that 1.58 was used in conjunction with 18.3 m^2 in Fw documents? Or are we just assuming this number 18.3 for some reason?

All lift and drag coefficients in that table refer to 18.3m². I don't think it's evident from the table alone, but it is evident from Fw material elsewhere. 18.3 is the reference area these coefficients refer to.

November 1, 2016

Exactly. The "error" is certainly not in the math. The "error" is that one creates a good source of mixing up the published CLmax for a profile with the maximum lift coefficient an airfoil can have. Those "CLmax'ens" have not much to do with each other. The "theoretical" one says when the profile achives maximum lift. The other tells you the starting point where a real world airfoil, consisting of various, different profile sections, has achived maximum lift. The numbers therefore cannot be equal. Them not being equal is also not dependent on a suposedly defective finish.

And yes, just one report, no matter from where it is, is nowhere near enough substance to overthrow current practises in aero engineering. Nature (welcome to the club, biologist ) has published some papers that "were not that good" as well.

Taken together, I stand by my original assessment that your spreadsheet is very useful nevertheless. One just has to be aware of the limits or what it tells you and what it doesn't. that's all.

And besides, the link I gave you shows an example of a software doing for routine what we are trying to reproduce here. That program is an elaborate and more detailed version of your spread sheet and it is just one of several software tools to do such.

None of them have a parameter "for the practical deviations of real world profile sections compared to published values". It is just not a topic at all. Nowhere. Therefore, I still must assume that published values are (like the ominous 1.58) used. So the NACA report was not considered important enough (even for NASA) to "include it" in current design principles. So, why should I?

I am beginning to think that you are deliberately misrepresenting what I have been saying. So I am just going to sum up my take on all of this one more time, because I would much rather that was not the case.

No-one is trying to overthrow established practices or look for "the practical deviations of real world profile sections compared to current values". BTW you really should not use quotes like that when that is not something that I ever said. I am not interested in the deviations of profile sections and neither is the calculator - it only uses the whole aircraft characteristics.

Why should you (as someone who wants to start from the airfoil data) care about the difference between profile, wing and aeroplane data when assessing the Fw's aeroplane CLmax? Because:

1) You have not "done the math" to generate an aircraft CLmax from the airfoil and other data in respect of the Fw190

2) I very much doubt that you could, even if you tried, since you do not know the effects of gunports, wheel wells etc on the lift generated

3) FW were not using modern computer aided iterative design systems either, no doubt they used some rules of thumb and experience and then checked empirically

4) The contemporary NACA reports, who would have been using similar methodology the Fw, are crystal clear in their conclusions. These effects matter. Ignore them if you like, but I find it hard to understand why anyone else should.

* * *

As to the formula/calculator: when I first posted it I emphasized that one had to be careful and that using it demonstrated the uncertainty due to the sensitivities, not what the right answer is, and have repeated this caution since.

Given the uncertainty about the actual empirical inputs no-one can be sure of any derived number's. The maths is indisputable but we just do not have the facts, consequently we cannot use that calculation to show "good agreement" with 1.58 or anything else.

What is possible is a range of outcomes consistent with a range of probable inputs. 1.58 is a member of possible outputs, but given what I have seen of the data, not a very likely one.

I have not yet seen a single cogent argument against this conclusion.

Lets look again at the variables in the context that the calculator was designed to explore, the RAE test:

1) CLmax is for the aircraft - airfoil is irrelevant. We have this number 1.58 from Fw documents - but no definitive statement of what it refers to. Probably the wing - but see 4 below.

2) Weight - we do not know - but, the US report on the RAE trials said without ammo

3) Vmin - we are not sure even what the IAS was: 110mph or 118mph, let alone what the PEC should be, given what I take to be two different reports about the same tests.

4) S - wing area. I used the 18.3 reference area since this was Crump's number. But what is the 15.6m^2 in the Fw data sheet? Does it relate to the 1.58 shown lower down? No-one is saying.

edit - just seen JtD's post above. Thanks for clarification - still wondering what the 15.6 is though it is probably irrelevant to this discussion.

5) rho - no idea so everyone is just assuming sea level standard day, and assuming that the speed was somehow corrected for that. Was it - no idea.

Just a bundle of assumptions - interesting to play with, but without much more tightly specified data, little can be concluded about real world outcomes except perhaps for the extreme limits of the possible range and - at best - some idea of what is more or less probable.

* * *

Trying to make out that it is all a misunderstanding and that Crump and Holzauge/JtD are really agreeing is incorrect. Some disputes do turn out like that: not this one.

Suppose we had a Fw190 in something like factory fresh condition, access to a wind tunnel and then we tested for Clmax. This is possible, we just cannot afford it.

Crump is saying that the result would be very close to 1.58: although how close is not really clear as you can see from his incoherent posts.

JtD/Holtzauge are saying that the result would probably be between 1.35 - 1.45, maybe a little more either side. IIRC.

They are making different predictions about an empirical test, they cannot both be right.

Edited November 1, 2016 by unreasonable

November 1, 2016

Unreasonable, I see that I failed making my point clear to you. It is also not my intentions to talk you down in any way. Sorry if I did and I certainly don't want proceed with an unproductive argument.

Btw., did you check out the link I posted for the software? Given you created a CLmax calculator, I thought you might find this interesting?

November 1, 2016

4) S - wing area. I used the 18.3 reference area since this was Crump's number. But what is the 15.6m^2 in the Fw data sheet? Does it relate to the 1.58 shown lower down? No-one is saying.

Crump is saying that the result would be very close to 1.58: although how close is not really clear as you can see from his incoherent posts.

It is 1.58 for the clean wing CLmax as measured thru comparative studies by the Focke Wulf Design team, confirmed by a modern VSAERO CFD analysis, and independent research by Grumman for their design with the same airfoil selection.

Pretty solid evidence.

You keep confusing the fact the 2D data CLmax is 1.7 and you just do not understand how that becomes 1.58 as our 2D data and gives good agreement with the aircraft.

You should probably read that sheet I posted on wing design, it might clear some things up for you.

One thing you can hold as a foundation of aeronautical science....the 2D data will agree with the airplane. It is the only way we can design airplanes to do specific jobs.

None of them have a parameter "for the practical deviations of real world profile sections compared to published values". It is just not a topic at all. Nowhere. Therefore, I still must assume that published values are (like the ominous 1.58) used. So the NACA report was not considered important enough (even for NASA) to "include it" in current design principles. So, why should I?

That NACA report is actually telling a designer what to troubleshoot if he does not get good agreement.

November 2, 2016

Unreasonable, I see that I failed making my point clear to you. It is also not my intentions to talk you down in any way. Sorry if I did and I certainly don't want proceed with an unproductive argument.

Btw., did you check out the link I posted for the software? Given you created a CLmax calculator, I thought you might find this interesting?

No problem - no doubt grumpy old man syndrome on my part

"For the Fallen" used to bring a tear to my eye thinking about the lost young soldiers - now it brings a tear to my eye thinking about the decrepit old gits.

They shall grow not old, as we that are left grow old:

Age shall not weary them, nor the years condemn.

* * *

I did look at the linked page, and it certainly looks like a neat piece of kit. It deduces aeroplane performance from physical first principles rather than having certain performance limits plugged in as in a FS FM, IIUC.

But I do not intend to buy it - I do not have X-plane to make the cheap version useful, and while I would happily slash the gf's allowance to buy the Pro version, I very much doubt that I would be able to make head or tail of it.

It does not surprise me that a modern PC tool like this can get fairly close to predicting RL data from modern aircraft that are well documented and also generally built using the various tricks that have been learned to reduce the gap between theoretical and empirical outcomes that seems to be the cause of much of the bother above. I do wonder how it would model a Fw190 - or even better, a Sopwith Camel

But you would still have to test the predicted/modeled results against a real plane to see if it had captured the relevant issues.

(Then even if you did test it you would have to deal with the "it is not a valid sample size" arguments. )

November 2, 2016

Ah, unreasonable, it's not about being a grumpy old man, it's about what you make of being one.

Some do have a precise idea of what to make from it

Same goes with the nature of a discussion in a forum. Telling from personal experience (as I do this for a living), properly understanding what a boffin is actually talking about and means to say is more often than not harder than the science he's working on.

Btw, unreasonable, I was wondering whether you have XPlane. I don't; but looking at the sofware package I linked above, I'm tempted to do so. I'm curious what could attain with using that and how the result would look like. Would be cool being able to making reasonable reproductions of know, "undisputed" FM models. Then see how it works when you tackle "disputed ones".

November 2, 2016

Proper understanding may be hard, but step one is to read the post.

But I do not intend to buy it - I do not have X-plane to make the cheap version useful, and while I would happily slash the gf's allowance to buy the Pro version, I very much doubt that I would be able to make head or tail of it.

November 2, 2016

I do not have X-plane

Indeed.

I used past tense because I was referring to myself wondering at the time when I was originally posting the link of the software. Did past tense imply a "present" matter? It was also the main reason why I put a link to this particular software in posting #221. Had I referred to your post #232, I had been wondering (referring to myself supposedly wondering, present tense, in posting #233). Correct? You would know.

How complicated. Too complicated.

See, communication is truly no trivial trivial matter and I for sure agree that in forums one tends to over read such subtleties. At least I do often enough, that is true.

Syntax difficult is and sources for misunderstanding plentiful are. Then threads locked get.

I guess we can move on now, leaving these subtleties. I'd be happy to.

November 2, 2016

I am not a linguist but..... - it is past tense but there are several forms. http://www.verbix.com/webverbix/English/think.html

So meaning here depends on context: in telling a story you could say 'I was walking alone on the beach deciding whether to end it all" and that clearly does not imply that you still are walking now, since the context situates the walking in a story about the past. But the form "was walking" does imply continuity up to some reference time. If the reference time is not clearly located in the past by the context, the implication is that the action is still continuing.

So "I was wondering whether X." without context definitely means that you are wondering at the time of speaking/writing. If you ask a native English speaker "I was wondering if you have a match" he will say "Sure" and light your cigarette, not ask you "Are still wondering or have you stopped?": unless he is being facetious. Some people - not me - think this form is more polite than a direct question "Do you have a match?" and it is often used in this way: that is how I took your remark.

Tricky if English is your first language, even worse if not.

BTW, where I live the language - Thai - has no conjugations at all, tense is indicated by simple prefixes and the verb never changes. Much easier!

This is all gloriously Off Topic, but that probably just reflects the fact that the topic is just about done for unless someone can come up with any relevant data.

November 3, 2016

Well gee, Crump doesn't even have the good grace or decency to accept a withdrawal and apology, without feeling the need to be condescending. Apology withdrawn.

November 3, 2016

I am getting sick and tired of cleaning up these 190 threads and editing out all the childish condescending smack talk and back and forth nonsense.

If this kind of stuff continues I will lock every single 190 thread that pops up in here.. and/or start to permanently ban certain forum members who always seem to be at the eye of every single FW 190 hurricane that has ever appeared anywhere..

Surely there is more than enough information in all these locked threads with all their charts graphs and techno babble to practically write a service manual for the 190 and to justify not allowing a single one to rear it;'s head for more than a few pages since they always seem to devolve into either insinuations or direct accusations of developer bias .. or back and forth pissing contests between a handful of key members.. You all know whop you are... so just knock it off.. UIn the meantime every locked 190 thread from here on will have the link below attached.. If that is not good enough then I don't lknow what to tell you. Build your own dadgummed sim and makle your 190 how you think it should be.

Fw 190 FM calims response

CLmax Calculator

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