CLmax Calculator

[ZachariasX]

After a boozy weekend this equals:

 

ClCl == nn WW // qq SS

 

Just in case Openoffice returns errors...

[CUJO_1970]

OK, so can a solid case for 1.35 with documentation be made then?

 

Or has it been made?

[=EXPEND=13SchwarzeHand]

It is currently being discussed. I guess Devs have other priorities right now though.

Edited October 22, 2016 by II/JG17_SchwarzeDreizehn

[unreasonable]

Well it seems to work OK, now that the lunch has worn off - amended calculator attached for someone to check my homework.

 

New calculation is underneath the old stuff, still all on one sheet. It takes default numbers from above, but you can write over anything in blue. You can either check for stall speed at given g, or test if Cl > Clmax  (assuming Clmax is a constant).

 

CLmax Calculator.zip

 

 

 

[NZTyphoon]

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.

 

In regard to the effects of wing armament, NACA published a Summary of Data Relating to the Effects of Wing Machine-gun and Cannon Installations on the Aerodynamic Characteristics of Airplanes .

 

In regard to the effects of wing surface drag on a P-47's wing: A Flight Investigation of the Effect of Surface Roughness on Wing Profile Drag with Transition Fixed

 

Another report that could be useful in the context of this thread; Effect of Mach and Reynolds Numbers on Maximum Lift Coefficient (1948): this report was based on flight tests  using the following: P-39N, P-63C, F6F-3, P-38F, YP-80A and P-51B.

[Holtzauge]

In regard to the effects of wing armament, NACA published a Summary of Data Relating to the Effects of Wing Machine-gun and Cannon Installations on the Aerodynamic Characteristics of Airplanes .

 

In regard to the effects of wing surface drag on a P-47's wing: A Flight Investigation of the Effect of Surface Roughness on Wing Profile Drag with Transition Fixed

 

Another report that could be useful in the context of this thread; Effect of Mach and Reynolds Numbers on Maximum Lift Coefficient (1948): this report was based on flight tests  using the following: P-39N, P-63C, F6F-3, P-38F, YP-80A and P-51B.

 

I already had the NACA TN1044 report on Mach number effects in my archives but I had forgotten about it and browsing through it it is interesting to see how much the Clmax is lowered at higher speeds. There are a number of such reports and one interesting aspect is that both the Me-109, Fw-190 and Spitfire had for their time quite good high speed characteristics. For the Spitfire which has an earlier generation airfoil the good high speed characteristics are mostly due to the thin wing, not the profile: A thicker wing is hit hard at compressibility. For the NACA 230-series, that the profile also had good compressibility characteristics was purely fortuitois and a bonus side effect from designing in the other characteristics since it was not a design concern at the time of their conception. The Me-109 2R1 profile it turned out was also good in this aspect due to its similarity to the NACA 230 series.

 

The report on the effects of the gun ports I have not seen before and was very interesting: For example, the delta Clmax on the P-51B is apparently as large as -0.12 which is huge, close to 10%. Also, other installations turned out to be better or worse than expected, at least for me: One other interesting thing was a cannon fairing installation that looked close to the Spitfire Hispano fairing which was quite good and had little impact. Another report to the archives so thanks for that!

 

Surface roughness is an issue but then I think it would be fair to assume about the same standard for all BoX planes? Even if aircraft are delivered in various conditions to different belligerents and units I guess we should assume that all virtual pilots have a dedicated crew chief who has sanded and polished up our aircraft right?

Edited October 23, 2016 by Holtzauge

[MiloMorai]

It is nice to see everyone getting along exchanging info without the lectures.

[unreasonable]

I think the BoX specification is "factory fresh" - which I take to be clean and without wear and tear, or repair faults, but also without any specific sanding, polishing, filling in gaps etc. 

 

Same for engines - run in but not worn. Makes sense to me given the decision choose a single standard.

 

Interestingly in RoF SP Career mode, every individual aircraft is slightly different. I suspect the team decided that this was an unnecessary coding complication that would only cause problems, if anyone noticed at all, so they ditched it for BoX.

 

What I am less sure about is whether actual factory delivered aircraft really did meet the standards that are estimated from the publisheds test results, but that is another matter.

[Holtzauge]

I think the BoX specification is "factory fresh" - which I take to be clean and without wear and tear, or repair faults, but also without any specific sanding, polishing, filling in gaps etc. 

 

Same for engines - run in but not worn. Makes sense to me given the decision choose a single standard.

 

Interestingly in RoF SP Career mode, every individual aircraft is slightly different. I suspect the team decided that this was an unnecessary coding complication that would only cause problems, if anyone noticed at all, so they ditched it for BoX.

 

What I am less sure about is whether actual factory delivered aircraft really did meet the standards that are estimated from the publisheds test results, but that is another matter.

 

OK, did not know that the developers had specified that but that makes sense and in addition fair since as NZTyphoon's reference shows, surface finish does make a difference but as long as its the same across the board that should be OK then. Same for engines. Also did not know that RoF had a variation in planes which seems like a nice added touch but tough luck if you start your campaign with a lemon then.....

[ZachariasX]

 

 

tough luck if you start your campaign with a lemon then.....

 

The lemon is usually behind the stick.

[Holtzauge]

The lemon is usually behind the stick.

True, but lemon x lemon = lemon**2 which is even worse!   

[ZachariasX]

True, but lemon x lemon = lemon**2 which is even worse! 

 

Like this it becomes: "[pick yours] bias!!!1!"

[Crump]

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.

 

Different airfoil.  Why is this even being advanced by "those who know"????

 

 

 NACA 24016 Airfoil cl=0.30 T=16.0% P=20.0%

1.000045 0.001679

0.998512 0.002009

0.993920 0.002992

0.986299 0.004614

0.975695 0.006849

0.962173 0.009661

0.945815 0.013011

0.926721 0.016850

0.905009 0.021124

0.880812 0.025779

0.854277 0.030754

0.825568 0.035991

0.794861 0.041426

0.762344 0.046997

0.728218 0.052639

0.692691 0.058285

0.655981 0.063866

0.618315 0.069310

0.579923 0.074544

0.541040 0.079488

0.501905 0.084063

0.462759 0.088188

0.423840 0.091783

0.385387 0.094771

0.347636 0.097078

0.310818 0.098639

0.275082 0.099396

0.240191 0.099183

0.206386 0.097735

0.174035 0.094867

0.143516 0.090490

0.115207 0.084622

0.089463 0.077378

0.066597 0.068960

0.046863 0.059629

0.030448 0.049682

0.017468 0.039421

0.007971 0.029124

0.001944 0.019022

-0.000677 0.009283

0.000000 0.000000

0.003760 -0.008510

0.010368 -0.015988

0.019659 -0.022514

0.031476 -0.028192

0.045673 -0.033143

0.062131 -0.037489

0.080763 -0.041348

0.101520 -0.044825

0.124387 -0.048004

0.149377 -0.050944

0.176517 -0.053671

0.205828 -0.056168

0.237311 -0.058369

0.270927 -0.060145

0.306499 -0.061303

0.343347 -0.061731

0.381168 -0.061464

0.419726 -0.060556

0.458782 -0.059067

0.498095 -0.057060

0.537419 -0.054603

0.576512 -0.051765

0.615131 -0.048611

0.653036 -0.045207

0.689993 -0.041615

0.725773 -0.037895

0.760154 -0.034103

0.792924 -0.030295

0.823880 -0.026525

0.852830 -0.022845

0.879594 -0.019309

0.904008 -0.015967

0.925919 -0.012870

0.945192 -0.010068

0.961707 -0.007606

0.975361 -0.005527

0.986070 -0.003868

0.993768 -0.002660

0.998406 -0.001926

0.999955 -0.001679

 

 

 

 

CA 23015 Airfoil cl=0.30 T=15.0% P=15.0%

1.000035 0.001575

0.998499 0.001879

0.993903 0.002786

0.986273 0.004282

0.975656 0.006342

0.962119 0.008936

0.945742 0.012023

0.926628 0.015560

0.904893 0.019497

0.880670 0.023783

0.854108 0.028362

0.825371 0.033179

0.794635 0.038175

0.762089 0.043293

0.727933 0.048472

0.692376 0.053650

0.655638 0.058764

0.617944 0.063747

0.579525 0.068529

0.540618 0.073038

0.501461 0.077201

0.462295 0.080942

0.423360 0.084187

0.384895 0.086864

0.347136 0.088904

0.310314 0.090249

0.274656 0.090847

0.240382 0.090658

0.207700 0.089657

0.176402 0.087787

0.146249 0.084735

0.117774 0.080218

0.091513 0.074140

0.067964 0.066589

0.047541 0.057813

0.030555 0.048172

0.017193 0.038075

0.007526 0.027921

0.001522 0.018048

-0.000930 0.008700

0.000000 0.000000

0.004012 -0.007767

0.010790 -0.014413

0.020104 -0.020094

0.031750 -0.025000

0.045565 -0.029337

0.061452 -0.033297

0.079396 -0.037046

0.099470 -0.040707

0.121820 -0.044351

0.146645 -0.047972

0.174150 -0.051468

0.204515 -0.054593

0.237120 -0.057036

0.271353 -0.058737

0.307002 -0.059714

0.343847 -0.059996

0.381660 -0.059624

0.420206 -0.058648

0.459246 -0.057126

0.498539 -0.055117

0.537841 -0.052687

0.576910 -0.049900

0.615502 -0.046818

0.653379 -0.043505

0.690307 -0.040017

0.726058 -0.036414

0.760410 -0.032748

0.793150 -0.029072

0.824077 -0.025437

0.852998 -0.021894

0.879736 -0.018492

0.904124 -0.015280

0.926012 -0.012306

0.945264 -0.009616

0.961761 -0.007255

0.975400 -0.005261

0.986097 -0.003671

0.993786 -0.002514

0.998418 -0.001811

0.999965 -0.001575

 

 

 

 

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

 

 

You are absolutely correct.  Kurt Tank and his team of engineers most certainly did know the CLmax of their design just as every other design team then and today does.

 

 

Reynolds number is not "The Undiscovered Country" in the 1940's.  It was well established design criteria that is still in use today in the exact same way.

 

Reynolds published his finding and proved them experimentally in the 1880's.

 

 

The subject of ratios and the opinions advanced by some is baffling.  

 

Anyway, if you understand Reynolds Number and what it means then it becomes a certainty that at the same Reynolds Number an airfoil MUST deliver the same Coefficient of Lift.

 

Reynolds Number is a non-dimensional ratio that represents the "stickiness" of the air.  At its basic level, Re number describes the flow characteristics in a non-dimensional ratio form and is a function of the length of object being examined.

 

What it tells us is that if we have a 61cm chord (long) airfoil under a specific "stickiness" of the air then it will achieve a set ratio of lifting pressure to dynamic pressure.

 

If we put a 181cm chord (long) airfoil under the same "stickiness" of the air then it will achieve the same set ratio of lifting pressure to dynamic pressure.

 

It is a given and fact.

 

It is also a fact that Aeronautical engineering is predicated on the ability to accurately predict the characteristics of the aircraft based off the airfoil selection.

 

Once more, calculations NEVER supercede or take precedence over measured data.  The ONLY calculation system that is on par with measured results is Computation Fluid Dynamics.  

 

 

 

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……..

 

 

 

The idea that any designer would simply accept airframe leakage around ailerons, root fairing, an weapon installations is simply not true.  99% of this is errors in manufacturing and easily eliminated at the construction jig.  In fact, that is whole point of many of these studies is to confirm that basic agreement relationship between 2D data, model, full scale, and flight testing.  Non-agreement is not some "Oh well...gee guys we tried..."  It is indicator of a solvable problem.

Edited October 23, 2016 by Crump

[Crump]

 

 

The subject of ratios and the opinions advanced by some is baffling.     Anyway, if you understand Reynolds Number and what it means then it becomes a certainty that at the same Reynolds Number an airfoil MUST deliver the same Coefficient of Lift.   Reynolds Number is a non-dimensional ratio that represents the "stickiness" of the air.  At its basic level, Re number describes the flow characteristics in a non-dimensional ratio form.   What it tells us is that if we have a 61cm airfoil under a specific "stickiness" of the air then it will achieve a set ratio of lifting pressure to dynamic pressure.   If we put a 181cm airfoil under the same "stickiness" of the air then it will achieve the same set ratio of lifting pressure to dynamic pressure.   It is a given and fact.   It is also a fact that Aeronautical engineering is predicated of the ability to accurately predict the characteristics of the aircraft based off the airfoil selection.   Once more, calculations NEVER supercede or take precedence over measured data.  The ONLY calculation system that is on par with measured results is Computation Fluid Dynamics.  

 

The relationship of lifting pressure to dynamic pressure at a specific "stickiness" of the air is fixed and does not change.

 

The ability to accurately express airspeed was not fixed in the 1940's.  End of Story. 

[JG13_opcode]

Different airfoil.  Why is this even being advanced by "those who know"????

I thought it was a typo as somebody mentioned.

 

Anyway, if you understand Reynolds Number and what it means then it becomes a certainty that at the same Reynolds Number an airfoil MUST deliver the same Coefficient of Lift.

I don't see anyone disputing this. However the sectional polar is not valid for finite wings.

Edited October 23, 2016 by JG13_opcode

[Crump]

 

 

However the sectional polar is not valid for finite wings.

 

It most certainly is and is the basis of all aerodynamic prediction.

 

 When we derive the 3D wing, it MUST meet the 2D data.  End of story.

[Crump]

When you derive the 3D wing, the Coefficient of Lift is solely based on the 2D wing data polar under the same Reynolds Number.

 

The actual angle of attack is changed as per induced angle but that does not affect the 2D data coefficient.

 

2D section Angle of Attack (Coefficient of Lift) = Actual Angle of Attack to the Relative Wind - Induced Angle of Attack

 

Rearranging and solving for Actual Angle of Attack:

 

Actual Angle of Attack to the Relative Wind = 2D section Angle of Attack (Coefficient of Lift) + Induced Angle of Attack

 

That is the basic principle and exactly why the NACA airfoil data is measured, compiled, and made available to engineers.

 

[Crump]

 

 

I thought it was a typo as somebody mentioned.

 

Understandable.  It is not a typo.  The general principle that report tells us is that mach number and Clmax have a relationship such that Reynolds Number and Mach number must be the same for our 2D data to conform with our 3D wing/airplane.  That is why at high altitudes, your stall speed increases until you reach the coffin corner. 

[Crump]

[JG13_opcode]

Go back and re-read what I wrote. Your response addresses something other than what I am talking about, so I think you are wasting your time.

[Crump]

 

 

Go back and re-read what I wrote. Your response addresses something other than what I am talking about, so I think you are wasting your time.

 

Ok, my response address the response that the 2D polar such as found in "Summary of Airfoil Data" is not applicable for a finite wing.  That is kind of exactly why that report was published so that engineers could relate measured infinite airfoil data to the finite wing.

 

 

 

However the sectional polar is not valid for finite wings.

 

[JG13_opcode]

At the tip of any real wing, lift goes to zero, irrespective of airfoil.

 

If you're implying that the coefficient at the tip of the wing is 1.5 or whatever the peak of the 23009 polar is, I'm afraid you are incorrect.

Edited October 24, 2016 by JG13_opcode

[ZachariasX]

If you're implying that the coefficient at the tip of the wing is 1.5 or whatever the peak of the 23009 polar is, I'm afraid you are incorrect.

 

I don't really understand the argument as such. But more like the whole wing/airfoil should have a corresponding CLmax as the published "infinite" 2D wing. If what I say would be the case, then the question arises:

 

"What is the minimal wingspan for a given airfoil to produce lift corresponding (sufficiently) to 2D data?"

 

For example, if I have a given profile and a given Reynolds Nr., and say, 1 meter chord, would 1 meter span of the airfoil be enough? Or 100 meters? At what length could I say, ok, now it should be in the close range of 2D data, given that I have super-duper surface finish?

 

Yes, I can increase lift of the wingtips by adding winglets etc., thus reducing the effect. But on a common airfoil, how is that handled by people who get paid to do such?

 

Another question comes to my mind, going through this thread:

 

Given unreasonable made this spreadsheet to make an educated guess on a plane's (max.)lift, if we are provided loadout and airspeed. In the context of:

then it becomes a certainty that at the same Reynolds Number an airfoil MUST deliver the same Coefficient of Lift.

 

How much variation can we expect in the Reynolds Number, given an aircraft given the aircraft under scrutiny is operating in some atmosphere with some density and some humidity and some temperature?

 

Anyone knows? I am not a subscriber to that kind of magazines from Wiley or Blackwell, so I don't have access. I guess it would be instructive info to add to unreasonables spreadsheet.

[unreasonable]

But more like the whole wing/airfoil should have a corresponding CLmax as the published "infinite" 2D wing. 

 

 

 

I am not sure why you think it should, I thought that even Crump agreed that it did not, after posting a text book explaining how a significant difference would arise. Hence his argument that a 2D airfoil of 1.70 would/could lead to a 3D wing of 1.58.  (Perhaps he has changed his mind in more recent posts I cannot see).

 

Anyway, the key differences are those between a 3D wing, 3D model and actual aeroplane.  This is why I think that his constant banging on about the airfoil or even wing Clmax is missing the point and obscuring the important issues.

Instead of reading the clear and literal meaning of the NACA tests on these differences, we get into an untestable and hypothetical discussion about the intentions of people long dead,as though that could prove some empirical fact about the Fw190's Clmax one way or another.

 

If I may be pedantic for a moment, I did not make the calculator to make an educated guess about the plane's maximum lift - or anything else, for that matter.  If the formula is correct, the aeroplane Cl max is correct, for a given set of inputs (and the airfoil or wing Clmax is strictly speaking, irrelevant). If there is any educated (or even uneducated ) guessing it is in the inputs, whichever variable you solve for. The calculator is just an interactive way for people to see the sensitivity of the results to changes in the inputs. So it works just as well if the CLmax is an input - as in Crump's initial worked example on the RAE test, and the Vmin is calculated.

 

On the last point - not sure about the relevence of the Reynolds number TBH - but the air density is an input number so you can make any change you wish  - or you could rearrange and solve for it if you really want!

[ZachariasX]

If I may be pedantic for a moment,

 

 

If I overinterpreted you, then I dodn^t meant to. Still I'm happy having your spreadsheet as I can use it in the way I mentioned. And it for sure it is instructive as it illustrates the relations of the variables in a more practical way than with the HP-35 pocket calculator back then.

 

Still I am surprised to see the deviations of ~0.2 for airfoils from "real world values" on an airfoil compared to what I thought were measured values from built airfoils in a wind tunnel as well. In the article you posted these were differences seen in a front line combat aircraft of the 40's. But to say that you cannot get closer to published values...? Or can I?

 

Why do you think one would publish data if what you get is so far off?

 

 

Imagine on the Titanic, on published numbers and "the reality":

 

"We are sinking! We are thinking!!"

"What are you thinking about?"

"I'm thinking about we never fit all on those few rafts! I previously read the capacity per raft, and we have space for 1'100 passengers, half of what is aboard!"

"Trust me, in cases of emergency you never attain these values. You cuddle up. There is space."

"So, then, where is my space?"

"Gone with the last raft, all taken by first class passengers. Keep in mind, they might feel cozy a fair bit earlier than you. Again, you never attain published values."

"OMG!"

"Relax and take a life vest. We do have one for everyone."

"But it's effin' cold in the water!"

"Relax, it's only 3 miles to reach solid ground. You'll make it!"

"Oh, great. It's night, can you point me in the direction I have to swim for that?"

"Downwards."

 

 

Z

[JtD]

"What is the minimal wingspan for a given airfoil to produce lift corresponding (sufficiently) to 2D data?"

Well, an aspect ratio of 6 (span/chord) will get you quite close already. Below 6 the effects get large, and at really low aspect ratios can't even use the same lift theories. However, to exactly reproduce 2D you do need infinite. Additionally, wing shape matters (elliptical is good), as do a lot of other design features like change of profile along the wing or washout.

 

How much variation can we expect in the Reynolds Number, given an aircraft given the aircraft under scrutiny is operating in some atmosphere with some density and some humidity and some temperature?

Here's a calculator.It lets you change viscosity. But as you can see, the biggest input is the speed the aircraft flies at, if you go at 0.5 g stall speed you'll have a completely different RN than at 5g stall speed.

 

Still I am surprised to see the deviations of ~0.2 for airfoils from "real world values" on an airfoil compared to what I thought were measured values from built airfoils in a wind tunnel as well. In the article you posted these were differences seen in a front line combat aircraft of the 40's. But to say that you cannot get closer to published values...? Or can I?

The airfoils built for the wind tunnel were extremely highly polished test subjects, they were meant to give theoretical data and were treated as such. They were extremely smooth, polished and clean. After installing these airfoils in a wind tunnel, someone went in with a piece of cloth and wiped off finger prints, because these would effect performance. Outside of the laboratory, these conditions were simply not achievable, neither were real wings as smooth, nor as clean.

Which is why some testing was also conducted with less than perfect airfoil conditions, and, depending on the condition, you'd get maximum lift coefficients down to 60% -70% of that of the ideal air foil at the same Reynold number.

[unreasonable]

 

Still I am surprised to see the deviations of ~0.2 for airfoils from "real world values" on an airfoil compared to what I thought were measured values from built airfoils in a wind tunnel as well. In the article you posted these were differences seen in a front line combat aircraft of the 40's. But to say that you cannot get closer to published values...? Or can I?

 

Why do you think one would publish data if what you get is so far off?

 

 

You can use the calculator like that, and I am glad you think it is interesting or handy. Just be careful.....

 

But I am still not sure why this is causing a problem - the deviations mentioned in the NACA are not from different tests of the same airfoils. The deviation is between the airfoil number or the wing number - which could be completely accurate as determined by tests - and the CLmax of a real physical aeroplane, as determined by tests.

 

Given that the aeroplane has a fuselage and other bits (to use the technical term) plus manufacturing tolerances and the effects of gun ports etc which are not captured in the airfoil numbers at all,  it would be absolutely astonishing if the Clmax for a real aeroplane was = the Clmax of the airfoil on which it's wing was based. The wings are the part that does the flying, the "bits" do everything else.

 

So then the question is how much is the difference. This seems to be a purely empirical matter, addressed by various posts so far. 

 

As for the Titanic analogy - the only thing that has hit an iceberg here is the theory that the Fw190 has a Clmax of 1.58....

[JG13_opcode]

I believe Crump is looking for some primary source docs he has, showing what CLmax it was actually able to attain.

[CUJO_1970]

So...the difference between Clmax of 1.58 and ~1.35 is the "other bits" - meaning the area of the fuselage and whether or not those should be factored into the calculation?

 

1.58 includes fuselage, etc?

 

1.35 does not? Only includes lift factor of wing only?

[unreasonable]

So...the difference between Clmax of 1.58 and ~1.35 is the "other bits" - meaning the area of the fuselage and whether or not those should be factored into the calculation?

 

1.58 includes fuselage, etc?

 

1.35 does not? Only includes lift factor of wing only?

 

My understanding of the difference of opinion based on documentation seen so far:

 

Crump believes that 1.58 is the Clmax of the wing, and of the 3D virtual model, and of the actual plane. Any observed differences are problems that will go away with proper maintenance.

 

JtD, Holtzauge believe that 1.58 is the Clmax of the wing, the 3D virtual model is somewhat less than this and the real plane less than that - but with little precision as to a "typical" number, hence 1.35 or 1.40 is a reasonable estimate. The differences are at least partly due to unavoidable facts to do with the characteristics of the non-wing parts, the finish of the wing and things like weapon installations that are not problems but features.

 

Personally I find the second view more convincing both from the evidence and arguments presented.

 

The calculator can only apply to the whole thing that is being measured - ie the whole real plane in the default assumptions.

 

(A question to the experts: I think that in theory you can apply it (formula or calculator) to any part of the plane with a lifting surface, and hence analyse wings and fuselage separately, but there might still be a problem in combining results due to the interaction of the parts. Since I do not suppose anyone has put a wingless Fw190 fuselage in a wind tunnel and analysed it's Clmax, I am not supposing that this is in anyway useful btw, but if you did have a complete result for any two of wings - fuselage - plane you should be able to deduce the result for the other piece? Or does interaction make the results non additive?)

[JtD]

It would be more of an advanced aerodynamic calculation than an actual measurement. Probably with better results if you measured plane & wings and calculated the rest.

[Holtzauge]

Yes, and calculation and CFD modeling will only get you so far: CFD is quite good when you have mostly attached flow but when you start to get separated flows like at Clmax it starts getting difficult also for CFD and predicting the effects of all the small gaps, joints and leakage does not make it any simpler since this level of detail is usually not included on CFD models of complete aircraft.

 

So in the end its the flight testing that will give you the results. A good example that its not easy is stall characteristics: If it was as simple as selecting the right profile and applying some Clmax coockbook rules then you have to ask yourself why do some aircraft have good stall characteristics and some don't?

 

You can of course stand on the shoulders of the designers before you and use sound design practices that worked before to attempt to design in a high Clmax in combination with good control characteristics like stall warning and spin recovery but I believe you will never completely remove the element of chance and lady luck or else all aircraft would be benign, docile with a high Clmax because who would not design in all that if it was easy?.........

 

Anyone who tells you different is probably the type of character who Mark Twain had in mind when he said: "All you need in this life is ignorance and confidence, and then success is sure."

[unreasonable]

On reflection I do not think you could deduce the Clmax and Vmin for the fuselage even if you had all of the variables for the wing and the whole plane, since you still have to assume one of Clmax or Vmin to deduce the other, apologies for the red herring. (Pretty hard to test the results, too )

 

It seems to me that this discussion has reached a natural end, unless someone comes up with some new evidence. Probably best to highlight that in a new thread if it materializes. 

 

Lots of interesting links and good discussion, thanks again to everyone who made a constructive contribution. 

[ZachariasX]

 

 

I believe you will never completely remove the element of chance and lady luck

 

Sometimes you know. Roy Chadwick for example. He produced a stellar aircraft, then a lemon, then a stellar aircraft et. until he died in a lemon. So you even knew how his next plane was before he would even start to design it.

[MiloMorai]

Chadwick died in what could be called a lemon (more like it didn't live up to expectations) but it was not the aircraft's fault.

[Holtzauge]

Sometimes you know. Roy Chadwick for example. He produced a stellar aircraft, then a lemon, then a stellar aircraft et. until he died in a lemon. So you even knew how his next plane was before he would even start to design it.

 

So there is a pattern here: Simple: All u gotta do is scrap every second design!

[Crump]

Crump believes that 1.58 is the Clmax of the wing, and of the 3D virtual model, and of the actual plane.

 

Excuse me but Focke Wulf believes and published it in the their aerodynamic data.  VSAERO agrees with that assessment and returns the same result.  Grumman in a completely seperate and parallel development using the same airfoil selection returns the same results.

 

Not surprising as the root airfoil determines our wings CLmax.

 

 

 

 

o...the difference between Clmax of 1.58 and ~1.35 is the "other bits" - meaning the area of the fuselage and whether or not those should be factored into the calculation?   1.58 includes fuselage, etc?   1.35 does not? Only includes lift factor of wing only?

 

 

This is not correct and the first time I heard it over at DCS, it was a major pause moment.  [Edited]

 

The reality is the lift formula does not care where the lift comes from.  It accounts for everything.  It does not care if the lift force comes from the fuselage or the pilot flapping his arms.  

 

 

In our mathematically closed system, it we use the wing area as our reference area, we use the wing CLmax to determine performance.  That is how aircraft performance math is done.

 

Non-wing lifting surfaces are considered minor contributors.  They are important for behavioral math but have inconsequential effects on performance.

 

Here is a breakdown of typical fuselage lift contribution found in some of the World War II fighters in a maximum performance turn:

 

Bell P-39Q Airacobra: 

Wings: 88.5% 

Fuselage: 10.1% 

Horizontal Tail: .9% 

Vertical Tail: .5% 

 

Bell P-63 Kingcobra: 

Wings: 100% 

Fuselage: 8.5% 

Horizontal Tail: -9.1% 

Vertical Tail: .6% 

 

Supermarine Spitfire IX: 

Wings: 97.6% 

Fuselage: 11.9% 

Horizontal Tail: -9.5% 

Vertical Tail: 0% 

 

Focke-Wulf Fw 190A-8: 

Wings: 96.1% 

Fuselage: 7.2% 

Horizontal Tail: -4.6% 

Vertical Tail: 1.3% 

 

Focke-Wulf Fw 190D-9: 

Wings: 87.9% 

Fuselage: 13.3% 

Horizontal Tail: -2.4% 

Vertical Tail: 1.2%

 

P-51D: 

Wings: 92.9% 

Fuselage: 7.7% 

Horizontal Tail: -.7% 

Vertical Tail: .1%

 

Once more, it is a total red herring in terms of PERFORMANCE MATH.  Behavioral math it makes a difference.  If I want to calculate the trim change then I need to know what is producing force on what axis.  If I want to know what given amount of force on a given amount of mass will do....who cares how much trim change is required to move from point A to point B.

 

Add up all the lifting surface breakdowns...they all equal 100% of the lifting contribution of the aircraft.  That is accounted for in the lift formula!!!

 

Once again, it does not matter where the lifting force comes from and if we want a proper performance estimate using the wing area we must use the wing Clmax....just as Focke Wulf and Grumman did!!

 

You need to expand your diplomacy by not using this forum to bad mouth other developers. 

Edited November 1, 2016 by Bearcat

[Crump]

 

 

ut I am still not sure why this is causing a problem - the deviations mentioned in the NACA are not from different tests of the same airfoils. The deviation is between the airfoil number or the wing number - which could be completely accurate as determined by tests - and the CLmax of a real physical aeroplane, as determined by tests.

 

The differences are mainly due to expression of airspeed measurements.  There is not some huge difference between the wing, the model, the full scale wind tunnel, and the flying airplane.  It is a science and not a religion or philosophy.

 

There can be difference to poor finish and poor construction which is the main reason for conduction performance checks of assembly line aircraft.  The results tell the engineering team if everything is being assembled properly and if it is not, then the airframe goes back to the fitters and is not accepted for service.  

 

TSAGI had huge problem with this in trying to get furniture plants to build wooden aircraft structures.  It does happen but ALL aircraft customers and manufacturers practice quality control.  

 

The question is do you want to model your base aircraft as a poorly finished and constructed version or do you want it to represent a properly constructed and finished example when new?

 

The next question then becomes if you are going to lobby for a wing designed specifically for a high Clmax to have a lower Clmax then how is that going to effect your relative line up?  Are you going to lobby for lower Clmax airfoils to be increased so that the relative performance picture is now skewed to their favor?

Summary of Airfoil Data is not "highly polished" or some special finish.  Seriously,  What would be the point in that?

 

It is a summary of tested and measured airfoil data under achievable circumstances engineers can use to build airplanes.

[Crump]

 

 

So...the difference between Clmax of 1.58 and ~1.35 is the "other bits" - meaning the area of the fuselage and whether or not those should be factored into the calculation?

 

To be clear, the ~1.35 comes from a calculated estimate using airfoil theory.  It also only tells you the point at which flow separation begins and not the stall point.  It NEVER supersedes measured data or CFD analysis.

[Crump]

A considerable amount of airfoil data has been accumulated from tests in the Langley two-dimensional low-turbulence tunnels.

 

Data have also been obtained from tests both in other wind tunnels and in flight and include the effects of high-lift devices, surface irregularities, and interference.

 

Some data are also available on the effects of airfoil section on aileron characteristics. Although a large amount of these data has been published, the scattered nature of the data and the limited objectives of the reports have prevented adequate analysis and interpretation of the results.

 

The purpose of this report is to summarize these data and to correlate and interpret them insofar as possible.

 

 

 

Detail data necessary for the application of the airfoils to wing design are presented in supplementary figures placed at the end of the paper.

 

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930090976.pdf

 

It is not some crazy fantasy conditions or unreachable highly polished surface finishes, JtD.  It is just very useful data for predicting performance and designing an aircraft.   It represents a normal aircraft painted surface.

 

The most applicable section to wing design is the Aerodynamic Characteristics of the airfoils presented in the Summary of Airfoil Data.

 

Here is the aerodynamic data necessary for the application of the NACA 23015 airfoil to wing design....

 

 

It is the measured data that gives the designer the Clmax and the cruise coefficients one can expect to achieve in a properly constructed and finished airplane.  End of Story.