CLmax Calculator

[JG4_dingsda]

AFAICT the difficulty is that FW didn't actually do stall testing.

 

This report from the Flugtechnische Versuchsanstalt Prag, 10.04.1943, might be interesting:

 

"Abkippmessungen im Fluge mit einer einfacheren Messeinrichtung an verschiedenen Flugzeugen. - Vorläufiger Teilbericht 4: Abkippmessungen im Fluge am Flugzeug Fw 190"

 

Found in the Bundesarchiv, so this one should be in Freiburg as well [https://www.deutsche-digitale-bibliothek.de/item/LR4QKGDKKIFSEBPZUIZEJVE3ZLBFSO4F]

Edited October 19, 2016 by JG4_dingsda

[Crump]

Z you need to find the right F6F wind tunnel test document - it was quoted/pictured in one of the previous Fw threads. It is the one where an actual plane "in operational condition" IIRC, was tested and got a CLmax of 1.40 ?

 

Cannot remember the exact number. Then they covered the gun ports, filled obvious cracks etc and tested again and got 1.50 - compared to the wing value of 1.60

 

I am not talking about the document with the range of different wing models etc. I do not think it was a Japanese document.....

 

Clearly they thought the test was worthwhile, even though you obviously do not, so I would be interested in how that fits with your approach that the manufacturer is always right.

 

I will see if I can find it, but searching this forum is not so easy....

 

I got +6.4m^2 addition to the wing to get 1.17 BTW, using the calculator and starting assumptions delivering 1.58 Fortunately that is the only variable in the RAE test we can be reasonably sure of.

The 1.4 is an average of accelerated flight. Rate of change in angle of attack has a huge impact on CLmax. That is one of the biggest reasons you see such a large variation in calculated CLmax from flight tested stall speeds. That is the reason why all modern stall demonstrations list a speed loss as part of the standards.

 

It is also why Grumman used the NACA instrumentation to measure control surface deflection in flight.

 

JtD.....of course companies used other companies research tools. It happens all the time in aviation. It is a great way to make money renting or selling assets. The profit margins in aviation are small. If Focke Wulf needed to purchase wind tunnel time from a competitor they did it. Their competitor was more than happy to pocket the money and sell them the time. You have advanced a red herring argument. There where research facilities available for Focke Wulf to do their measurements and comparative studies.

 

JG13_OPCODE.....The measured 2D data CLmax at the Re the FW190 stalled at is 1.7. The EXTREME end of the normal loss for spanwise flow is 7%.

 

1.7-(1.7*.07)= 1.58

 

Anything outside of that normal range is cause to go back to the drawing board.

Edited October 19, 2016 by Crump

[Crump]

An example of typical behavior of companies sharing research assets is found in the famous Lockheed Skunkworks

 

All the major US aircraft manufacturers have offices there and use the facilities.

[ZachariasX]

JtD.....of course companies used other companies research tools. It happens all the time in aviation. It is a great way to make money renting or selling assets. The profit margins in aviation are small. If Focke Wulf needed to purchase wind tunnel time from a competitor they did it. Their competitor was more than happy to pocket the money and sell them the time. You have advanced a red herring argument. There where research facilities available for Focke Wulf to do their measurements and comparative studies.

 

 

---this is OT---

 

Nazi Germany operated slightly different than the Americans in your example. They basically never sold a product because it was cheaper in terms of dollars, but they sold it if they could argue along the lines that they could produce with "less effort" which translated into percieved economies of scale.

 

Heinrich Koppenberg was the main driver behind that, effectively turning Junkers into one of the largest Companies in the industry. It got that big through what can almost be called expropriation, turning formerly independent suppliers (with the know-how of manufacturing whole planes) into simple manufacturers of components. In battling Heinkel and Messerschmidt they almost distributed the whole industry amongst themselves.

 

The obsession of manufacturing few types but those at higher pace and numbers was the reason that killed most types, such as the Uhu for example. Eventually, they took his cockpit.

 

They didn’t have that much of large R&D because it simply didn’t pay. If all you produce are a handful of plane types, then you don’t plan for many and big installations such as wind tunnels for full size planes.

 

The R&D they had, like in Peenemünde etc. was more “common goods” made for specific weapons rather than an asset to a company that can maintain technological leadership.

 

---OT--

 

 

I am pleased that the calculator served as a tool to stimulate discussion - that is what I had hoped would happen.

 

 

Sure did!

[Holtzauge]

This report from the Flugtechnische Versuchsanstalt Prag, 10.04.1943, might be interesting:

 

"Abkippmessungen im Fluge mit einer einfacheren Messeinrichtung an verschiedenen Flugzeugen. - Vorläufiger Teilbericht 4: Abkippmessungen im Fluge am Flugzeug Fw 190"

 

Found in the Bundesarchiv, so this one should be in Freiburg as well [https://www.deutsche-digitale-bibliothek.de/item/LR4QKGDKKIFSEBPZUIZEJVE3ZLBFSO4F]

That certainly looks interesting judging by the title. Any chance you could get hold of it?   

[JG13_opcode]

 

 

JG13_OPCODE.....The measured 2D data CLmax at the Re the FW190 stalled at is 1.7. The EXTREME end of the normal loss for spanwise flow is 7%. 1.7-(1.7*.07)= 1.58 Anything outside of that normal range is cause to go back to the drawing board.

 

That's not what I'm asking.  What I'm asking is if you or someone else has primary source documents showing that the 190 achieved the design CLmax.

[Crump]

The RAE flight test of WNr 313 and Wright Patterson EB-104 both give good agreement with Focke Wulfs published CLmax.

[Crump]

---this is OT---

 

Nazi Germany operated slightly different than the Americans in your example. They basically never sold a product because it was cheaper in terms of dollars, but they sold it if they could argue along the lines that they could produce with "less effort" which translated into percieved economies of scale.

 

Heinrich Koppenberg was the main driver behind that, effectively turning Junkers into one of the largest Companies in the industry. It got that big through what can almost be called expropriation, turning formerly independent suppliers (with the know-how of manufacturing whole planes) into simple manufacturers of components. In battling Heinkel and Messerschmidt they almost distributed the whole industry amongst themselves.

 

The obsession of manufacturing few types but those at higher pace and numbers was the reason that killed most types, such as the Uhu for example. Eventually, they took his cockpit.

 

They didn’t have that much of large R&D because it simply didn’t pay. If all you produce are a handful of plane types, then you don’t plan for many and big installations such as wind tunnels for full size planes.

 

The R&D they had, like in Peenemünde etc. was more “common goods” made for specific weapons rather than an asset to a company that can maintain technological leadership.

 

---OT---

Yes, in Nazi Germany the the R&D facilities were state run/funded. Most of them in the United States are too including the Skunkworks. It is impossible to make a profit running an aerospace R&D facility. It does not eliminate a patent claim if your company was an innovator. Try putting wingtip fences on a design and not pay royalties to Airbus for example....if you do then your going to need a good legal team.

[JG13_opcode]

The RAE flight test of WNr 313 and Wright Patterson EB-104 both give good agreement with Focke Wulfs published CLmax.

 

Got a link or a pdf copy?  I can't find it on Mike's site.

[unreasonable]

Got a link or a pdf copy?  I can't find it on Mike's site.

 

Link to EB-104 is here http://www.wwiiaircraftperformance.org/fw190/eb-104.html

 

Interestingly this one does have a PEC correction graph! But the AIS is only shown down to 130-140 mph.  Reading through it, I can see nothing about a stall speed at all, (it is mentioned that "the stall is gentle")  so I am not sure how it can show "good agreement" even before addressing the issues of sensitivity that the calculator helped to expose.

 

The text of the RAE report which includes a stall test is here:  http://www.shockwaveproductions.com/store/fw190/tactical_trials.htm

 

If there were appendices to this with more data I have not managed to find them.  

 

The other detailed Fw test results from RAE were as per below, (from Crump's post) but these only address the rate of roll.

 

RAE 1231 part I.pdf

RAE 1231 Part II.pdf

RAE 1231 Part III.pdf

[JG13_opcode]

Yeah I've seen those.

 

I sincerely doubt the developers will accept 190G data, and the RAE test doesn't really give enough info.

[-TBC-AeroAce]

OMG what part of if u have the correct numbers Bernoulli;s cant be wrong do you not get!

 

And also why is CLmax so much of a fascination to you guys! 

 

Lift to weight vs power is so much more important!

 

CLmax is an easy number to get!!! I have worked on micro UAVs to the redesign of the A320! 

 

What does Clmax give u?

[unreasonable]

OMG what part of if u have the correct numbers Bernoulli;s cant be wrong do you not get!

 

And also why is CLmax so much of a fascination to you guys! 

 

Lift to weight vs power is so much more important!

 

CLmax is an easy number to get!!! I have worked on micro UAVs to the redesign of the A320! 

 

What does Clmax give u?

 

You can calculate an exact CLmax from a set of data, but the problem is that the numbers we have are insufficiently exact - changes in Vmin make larger changes in CLmax.

 

As to the second question as to why the CLmax obsession:

 

Relevance for the game is the critical AoA . Players are complaining that the critical AoA is too low and are getting into accelerated stalls too often.

I do not think anyone really cares about a few kph difference in the minimum stall speed, but a few degrees extra AoA would make a big difference to how the plane handles in a fight.

 

Also you know how the A-type personalities that tend to post on CFS forums just like to win an argument!

[JtD]

IIRC then RAE stated the Fw-190 wing areas as 177 sqft exposed and 197 sqft reference area. Going metric this equates to 16.44 and 18.3 sqm.

Good post, small addition here: Focke Wulf figures for the Fw190 are 15.6 and 18.3 respectively.

You will be getting additional lift from the horizontal stab, and, as you said, from the fuselage. As usual, things are a bit more complex than what can be put into a single paragraph.

 

In this sense, my assumption is that when you create a wing, not what is by convention also included between the outermost tips of a wing, then one should be able to reproduce measured lift qualities as published alson with the profile.

You need to really look at the cross sections of the wing as constructed - you will find that large parts of it have very little to do with the properties of the underlying ideal airfoil. In an extreme case, look at a cross section across the wheel well where wheel covers are partially removed. This alters aerodynamics considerably, not just in that cross section, but to some extend also to the left and right of it. The same is true for gun ports in the leading edge, or gaps at ailerons, imperfect sealing of hatches and so on.

On the bottom line, you won't be able to reproduce airfoil data with a manufactured wing. It won't produce the same high lift at the same low drag. It will resemble some of the airfoil properties, but to what extent, the guys back in the day could not be certain of.

 

I sincerely doubt the developers will accept 190G data, and the RAE test doesn't really give enough info

Actually, after removing the wing racks, there's little to chose between the A and the G. Personally I consider the EB104 test the best Allied tests of the Fw190 when it comes to detail and accuracy. I consider it fairly representative for A models as well.

[LLv34_Flanker]

S!

 

 Oh boy what you guys could do if working together instead of picking posts apart. After you filter the inevitable ego stuff etc out the information here is very interesting to someone like me who has no grasp of aerodynamics but the very basics.

[MiloMorai]

Yes, in Nazi Germany the the R&D facilities were state run/funded. Most of them in the United States are too including the Skunkworks. It is impossible to make a profit running an aerospace R&D facility. It does not eliminate a patent claim if your company was an innovator. Try putting wingtip fences on a design and not pay royalties to Airbus for example....if you do then your going to need a good legal team.

 

Learjet exhibited the prototype Learjet 28 at the 1977 National Business Aviation Association Convention. The Model 28 prototype employed the first winglets ever used on a jet and a production aircraft, either civilian or military. Does Boeing pay royalties to Learjet?

 

Skunk Works is an official alias for Lockheed Martin's Advanced Development Programs (ADP).

[Crump]

Got a link or a pdf copy?  I can't find it on Mike's site.

I can give you the data. I am at work so give me a day or two.

 

Unreasonable.....if you cannot correctly convert the speeds...you will not get the correct result.

 

As you have learned, a small change in velocity represents a large change in coefficient of lift at low velocity. Dynamic pressure is a function of velocity squared.

Learjet exhibited the prototype Learjet 28 at the 1977 National Business Aviation Association Convention. The Model 28 prototype employed the first winglets ever used on a jet and a production aircraft, either civilian or military. Does Boeing pay royalties to Learjet?

 

Skunk Works is an official alias for Lockheed Martin's Advanced Development Programs (ADP).

Winglets are not wingtip fences.

[Crump]

Airbus launched a legal battle on 2 December 2011. Anticipating a patent infringement lawsuit by Aviation Partners, Airbus filed a lawsuit to have the court declare the sharklet design was not copied from the Aviation Partners blended winglet.

 

Aviation Partners has escalated the patent dispute with Airbus by asking a US court in Seattle to slap an injunction on all new sales of sharklet-equipped Airbus A320s.

 

The injunction request by Aviation Partners on 31 July 2012 is the latest and potentially most damaging move in an eight-month dispute between Airbus and Aviation Partners over the intellectual origins of the A320 sharklet.

 

Aviation Partners accuses Airbus of “copying” the sharklet design using Aviation Partners’ proprietary information and data supplied to the airframer under a non-disclosure agreement, the Seattle-based supplier says in court documents.

 

Source of information for this section “Legal” is Flightglobal.

 

 

http://theflyingengineer.com/flightdeck/winglets-and-sharklets/

 

It is actually shaping up to be the largest legal battle in aviation history over patent infringement since Curtiss and the Wrights fought over the aileron.

[Crump]

Here is a quick primer on the winglets and the issues behind it.

 

http://www.dept.aoe.vt.edu/~mason/Mason_f/WingtipDevicesS04.pdf

 

 

Perhaps we can get back on topic and simply conclude the obvious. Research facilities were available for Focke Wulf to do comparative studies on the measured CLmax of their design.

[unreasonable]

Unreasonable.....if you cannot correctly convert the speeds...you will not get the correct result.

 

As you have learned, a small change in velocity represents a large change in coefficient of lift at low velocity. Dynamic pressure is a function of velocity squared.

 

 

For goodness sake, Crump, knock it off with the patronizing twaddle. I learned that a variable with a power would create a high sensitivity in physics classes at school, about 45 years ago, and have been dealing with this fact in my professional life in business and finance for the last 30 or so years.   How do think compound interest works?  You do not have to know anything at all about aerodynamics etc to see that this applies in this case - just look at the equation. 

 

What was amazing about all of your posts was that you never acknowledged that fact - I can only assume because you did not realize it.  If this had been so obvious to you, why did you not point it out, instead of claiming a totally specious accuracy in your RAE test "calculation".

 

Perhaps you did not realize this - so ignorance is the explanation.

Or you did realize it, but preferred to conceal the fact since it cast doubt on your results - so intellectual dishonesty.

 

Hard to tell - both are so plausible in your case.

 

Whatever, I have just had enough of your nonsense, time to go back on my ignore list. Adios! 

[ZachariasX]

On the bottom line, you won't be able to reproduce airfoil data with a manufactured wing. It won't produce the same high lift at the same low drag. It will resemble some of the airfoil properties, but to what extent, the guys back in the day could not be certain of.

 

Why would the publish a CLmax of standard airfoils to the accurancy of at least 1/100th, when for practical purposes that number always supposedly varied up to 1/3rd? Your telling me that Kurt Tank et al. had no precise idea about the lift that the airfoil of his design would yield (or only maybe in the range of 4 sqare meters wing give or take)?

 

What would be your guess on how the designers at that time started out designing a plane? What did they know? What didn't they know (of the things that they knew they didn't know)? I'm asking because I don't know, not having been there

 

Gunports, long portruding cannons, uncovered wheel wells, etc. for sure mess up the aerodynamic "cleanness" of your design. But do I understand you correctly, that you are saying one can't know about these effects reasonably well ex ante?

 

Given German planes (at least in the first 4 years of war) had a comparatively good surface finish, why should it suffer so much? Do I remember correctly that the Yak1 we have in game with its ClarkY has a higher CLmax, when the profile in theory delivers a lower CLmax than the NACA 230xx? Why would anyone use then the NACA 230xx profile series given that ClarkY is much easier to put in the jigs to build it? (It also has more benign stall qualities.)

[CUJO_1970]

This is getting depressing.

[Hoots]

This is getting depressing.

Getting?

[JG4_dingsda]

That certainly looks interesting judging by the title. Any chance you could get hold of it? 

I will not be able to go to Freiburg until second half of November -- I'll go there then, should nobody else have gotten his/her hands on it.

[Crump]

It is not my fault unreasonable if you did not understand the meaning and implication of why I so carefully converted the velocity for you in my very first post on the subject.

[MiloMorai]

It is not my fault unreasonable if you did not understand the meaning and implication of why I so carefully converted the velocity for you in my very first post on the subject.

 

"Garbage in = Garbage out" is careful conversion?

[Cloyd]

[snip] Whatever, I have just had enough of your nonsense, time to go back on my ignore list. Adios! 

Unreasonable, I continue to read your topic with fascination, without any understanding of what you guys are talking about. But I know enough to say that the "ignore list" is a beautiful thing.

 

Cloyd

Edited October 21, 2016 by Cloyd

[JtD]

Why would the publish a CLmax of standard airfoils to the accurancy of at least 1/100th, when for practical purposes that number always supposedly varied up to 1/3rd?

So that they could determine the relative performance of airfoils to a sufficiently accurate degree.

Your telling me that Kurt Tank et al. had no precise idea about the lift that the airfoil of his design would yield (or only maybe in the range of 4 sqare meters wing give or take)?

No precise idea about the lift of the wing, mainly. Airfoil fairly accurately, I'd say, for as long as the cross section of the wing actually matched the airfoil. They'd also know that they'd lose quite a bit of lift, so the uncertainty would not be in the 20+% range - it was certain it would not be anywhere near the theoretical airfoil maximum.

 

What would be your guess on how the designers at that time started out designing a plane? What did they know? What didn't they know (of the things that they knew they didn't know)? I'm asking because I don't know, not having been there

I haven't been there either, but generally they started with a target in the distance and their experience as a means to get there. Some folks maintained the wing and changed everything else, others changed the wing and maintained everything else when they put their next design on the drawing board.

 

The fact that this wasn't an exact science is evident from the comparatively large number of re-designs, design failures and accidents that occurred with quite a few of the new designs back in the day. I'm sure that you're also aware that neither performance, nor handling expectations were always met, even when everything looked great on the drawing board.

 

Gunports, long portruding cannons, uncovered wheel wells, etc. for sure mess up the aerodynamic "cleanness" of your design. But do I understand you correctly, that you are saying one can't know about these effects reasonably well ex ante?

That certainly depends on your definition of "reasonably". I'm stating that they had no means of accurately knowing the effect. Hence the necessity of confirming assumptions in testing.

Additionally, when guns were added as an afterthought, I'm sure the effects weren't considered in the early design phase when the airfoil was being chosen.

 

Given German planes (at least in the first 4 years of war) had a comparatively good surface finish, why should it suffer so much? Do I remember correctly that the Yak1 we have in game with its ClarkY has a higher CLmax, when the profile in theory delivers a lower CLmax than the NACA 230xx? Why would anyone use then the NACA 230xx profile series given that ClarkY is much easier to put in the jigs to build it? (It also has more benign stall qualities.)

I'm not sure I'm getting the question...but in case your wondering why the maximum lift coefficient of the Fw190 is lower than that of the Yak-1 in game - imho, there's no technical reason. Which is why it's been reported to the devs as implausible.

As to why anyone would use a NACA230xx over a ClarkYH - the NACA230xx offers a better lift-drag ratio as well as a better change of moment over angle of attack. At the expense of being harder to manufacture and having a more abrupt stall, as you've said.

 

Edit: No offence meant with this multi-quote thing.

Edited October 21, 2016 by JtD

[unreasonable]

Unreasonable, I continue to read your topic with fascination, without any understanding of what you guys are talking about. But I know enough to say that the "ignore list" is a beautiful thing.

 

Cloyd

 

It sure is, cheaper than beta-blockers and works better too! 

[Holtzauge]

I will not be able to go to Freiburg until second half of November -- I'll go there then, should nobody else have gotten his/her hands on it.

 

Great! Looking forward to seeing what it says!

[NZTyphoon]

For interest, here is a NACA report Effects of Compressibility on the Maximum Lift Characteristics and Spanwise Load Distribution of a 12-Foot-Span Fighter-Type Wing of NACA 230-Series Airfoil Sections dated November 1945: the model wing tested has a root section of NACA 23016 (sic! page 6: There was no NACA 23016 - this should read NACA 23015) and a tip section of NACA 23009...

 

Also, for interest is a much bigger report Summary of Airfoil data dated March 1945, which discusses data for most of the NACA airfoils developed to that time.

Edited October 21, 2016 by NZTyphoon

[JG13_opcode]

For interest, here is a NACA report Effects of Compressibility on the Maximum Lift Characteristics and Spanwise Load Distribution of a 12-Foot-Span Fighter-Type Wing of NACA 230-Series Airfoil Sections dated November 1945: the model wing tested has a root section of NACA 23016 (sic! page 6: There was no NACA 23016 - this should read NACA 23015) and a tip section of NACA 23009...

 

Also, for interest is a much bigger report Summary of Airfoil data dated March 1945, which discusses data for most of the NACA airfoils developed to that time.

 

Great find.  For the uninitiated, page 9 of the first report in particular goes into some detail on the multifaceted nature of and complications involved with aircraft design, and how numerous factors can come together to dictate the max obtainable lift coefficient (in their proposed example limited by tail effectiveness).  

 

A quick glance at Figures 7 and 8 within indicates a wing CLmax at the 190's stall speed (about 0.144 of mach at sea level in the 1976 std atmosphere) of about 1.35, which IMHO strongly supports the assertion made elsewhere of 1.3-1.4 for the 190.

 

Thanks for posting, 

Edited October 22, 2016 by JG13_opcode

[unreasonable]

For interest, here is a NACA report Effects of Compressibility on the Maximum Lift Characteristics and Spanwise Load Distribution of a 12-Foot-Span Fighter-Type Wing of NACA 230-Series Airfoil Sections dated November 1945: the model wing tested has a root section of NACA 23016 (sic! page 6: There was no NACA 23016 - this should read NACA 23015) and a tip section of NACA 23009...

 

Also, for interest is a much bigger report Summary of Airfoil data dated March 1945, which discusses data for most of the NACA airfoils developed to that time.

 

+1  for data - But there is a lot to absorb in those, especially for us non-specialists: is there anything in particular you would point out that we should pay attention to in the context of the discussion?  

 

(I see JG_13 opcode has already drawn attention to one point).

 

I also note from page 9. of the "Effects of Compressibility..." report " As the Mach number increases above 0.3 the angle of attack at which the wing stalls progressively decreases..."

 

I wonder if it does in BoX's FM?  If this phenomenon is captured in the FM this might also help explain why players are finding the accelerated stall so difficult: if you start with an critical AoA that is too low by a few degrees at Vmin, and then (correctly) model the reduction as speed increases, by the time you are at maximum speed in the M 0.5 region you have very little left.....

 

This is just thinking aloud.... 

Edited October 22, 2016 by unreasonable

[JG13_opcode]

+1  for data - But there is a lot to absorb in those, especially for us non-specialists: is there anything in particular you would point out that we should pay attention to in the context of the discussion?  

 

(I see JG_13 opcode has already drawn attention to one point).

 

I also note from page 9. of the "Effects of Compressibility..." report " As the Mach number increases above 0.3 the angle of attack at which the wing stalls progressively decreases..."

The Summary of Airfoil Data report is literally just a collection of polars for a variety of airfoils and doesn't tell us anything we don't already know. 

 

I wonder if it does in BoX's FM?  If this phenomenon is captured in the FM this might also help explain why players are finding the accelerated stall so difficult: if you start with an critical AoA that is too low by a few degrees at Vmin, and then (correctly) model the reduction as speed increases, by the time you are at maximum speed in the M 0.5 region you have very little left.....

 

This is just thinking aloud....

 

The report goes on to speculate that it's predominantly caused by Reynolds Number effects, rather than due to Mach. Either way, if the underlying physics of the sim are sound, then this will be captured by their implementation of the fundamental laws of fluid flow. I imagine the scenario you're describing would be well described by adding the load factor to your spreadsheet to illuminate just how fast you can exceed the CLmax by pulling G's.

Edited October 22, 2016 by JG13_opcode

[JtD]

Thanks for pointing out that report, hadn't seen it before.

 

A quick glance at Figures 7 and 8 within indicates a wing CLmax at the 190's stall speed (about 0.144 of mach at sea level in the 1976 std atmosphere) of about 1.35, which IMHO strongly supports the assertion made elsewhere of 1.3-1.4 for the 190.

The test configuration varies in two important points from the Fw190 wing, which pretty much cancel each other out. Firstly, the Reynolds number at Mach 0.14-0.15 is much lower in this test than it would be for the Fw190 (~2 instead of ~6*10^6). This costs ~0.2 in the maximum lift coefficient. Secondly, the wing is considerably smoother than the Fw190 wing, which gives ~0.2 in the maximum lift coefficient.

 

There are similar tests conducted with the same wing in the 19' tunnel, were two different densities (thus different Mach-Reynolds relation, see figure 4) are tested, as well as the wing with a rough leading edge (rough is pretty much icing, not production finish, but shows a trend). The above report is frequently used as a reference (5).

 

Technical Note 1299

[JG13_opcode]

Firstly, the Reynolds number at Mach 0.14-0.15 is much lower in this test than it would be for the Fw190 (~2 instead of ~6*10^6). This costs ~0.2 in the maximum lift coefficient. Secondly, the wing is considerably smoother than the Fw190 wing, which gives ~0.2 in the maximum lift coefficient.

Hmm, I agree about the low Re, but I skimmed the whole report earlier and didn't catch any mention of surface polish. I'll have to take a 2nd look. Thanks for the other link.

[unreasonable]

. I imagine the scenario you're describing would be well described by adding the load factor to your spreadsheet to illuminate just how fast you can exceed the CLmax by pulling G's.

 

That sounds horribly like a homework assignment!  Not really suitable for a boozy weekend.... perhaps I can get my head around it on Monday. Or not.... 

[ZachariasX]

 

 

Edit: No offence meant with this multi-quote thing.

 

Works perfectly for me. Thanks for the clarification of your point of view.

[Holtzauge]

I think there have been a lot of good input and references posted above on why it’s difficult to design and predict how a wing will actually perform but here are some added points:

 

In addition to it being difficult to predict how small deviations in the wing profile on a micro scale like joints, bulges, gun ports, hatches, rivets and surface finish etc. the IRL wing on a macro level will also deviate from a wind tunnel model in that there are gaps between ailerons, flaps and access hatches etc. that are not airtight and since air will flow from high to low pressure you will get small disturbing “fountains” of air seeping out from under and inside the wing to the upper suction side disturbing the flow. In a solid smooth wind tunnel model and a CFD model this is of course not an issue. You also have effects from propeller wash, wing root fillets etc. the effects all of which may be difficult to predict. So even if you can run at full scale Reynolds numbers (which was the exception not the rule) like in NACA’s variable density wind tunnel or get CFD results you still have to do some guesswork. So in the end, even if you apply sound design rules learned from earlier designs, a good wing profile and build in a slight washout this is still no guarantee you will end up with a low drag, high lift wing with benign stall characteristics. You really won’t know till you fly it IRL.

 

Regarding the NACA 230 series used on the Fw-190 this was and still is one of the best all round profiles ever developed and was the result of extensive trials in the NACA variable density wind tunnel (i.e. it and the others were evaluated at the right Re number) by the NACA scientist Eastman Jacobs who painstakingly tried out a multitude of different wing profiles of different thicknesses and camber distributions. He concluded IIRC that the best all round profile was the NACA 23012. The NACA 230-series was something new in that it had the highest point in the camber line more forward than was usual at the time. This is something akin to a leading edge flap which is also the explanation for the relatively high Clmax since the profile has a slight leading edge droop built in. In addition, this new forward loading profile meant that it had much better pitching moment characteristics than other more conventional profiles which had a more rear loading and is also why the Fw-190 does not need to retrim so much for different speeds. So all these good features were combined with decent drag characteristics which explains its popularity.

 

That the Fw-190 and most other designs had a profile in the order of 10% at the tip and around 15% at the root is due to the fact that if you go below ca 10%, the Clmax of many profiles will drop drastically. That the root is thicker is beneficial for two reasons: first of all the profile usually has to be thicker in order of 15% to be able to fully utilize the lift boost that a flap gives. A thinner profile will with flap deflection stall at a much max lower aoa. In addition, it is much easier for the structural engineers to design light weight wing spars as the deflection of a wing goes down with the cube of the wing section height. Also you need space for armament and landing gear. So there are plenty good reasons why the root usually has a higher thickness ratio than the tip.

 

So to sum up, making a good wing is not as trivial as choosing a profile simply based on its 2D aerodynamic characteristics and then applying some good design practices and presto you have a winner. In addition, Clmax is just one thing you need to keep in mind: High speed drag, pitching moment characteristics, the possibility to support full flap deflection in the inner sections and structural and payload concerns all have to be balanced. So in the end it’s a complex compromise just like most things in life are……..

[JG13_opcode]

That sounds horribly like a homework assignment! Not really suitable for a boozy weekend.... perhaps I can get my head around it on Monday. Or not....

Cl = n W / q S

 

Where n is the load factor, W is the weight, q is the dynamic pressure and S the reference area