unreasonable Posted October 15, 2016 Posted October 15, 2016 After following the Fw190 wars raging as usual I thought it might be helpful for people who are interested in the maths of how CLmax, weight and stall speed interact to have a little calculator. (Ie me, at least!) Not finding one, I built one, which I attach. It is done in SI units, but with conversions for imperial speeds and weights for those still unable to use scientific conventions. You can use it to calculate CLmax given other factors are known, including minimum stall speed. Alternately, you can calculate the minimum stall speed for a given CLmax and weight. Contrary to what some might have you believe, this is a fairly simple set of calculations. What is important is the sensitivities. Feel free to point out any errors. The spreadsheet is wrtten in Open Office, but it should open on your Micromonopolist program. CLmax Calculator.zip 3
SvAF/F16_Goblin Posted October 15, 2016 Posted October 15, 2016 Thank you, great for us illiterates that are not fluent in "aeronautics"
JG13_opcode Posted October 15, 2016 Posted October 15, 2016 (edited) Caveat emptor: That formula does not take into account things like compressibility/Mach effects. On many airfoils, as you get faster and small localized shockwaves start to form the maximum lift coefficient actually decreases. Edited October 15, 2016 by JG13_opcode
unreasonable Posted October 15, 2016 Author Posted October 15, 2016 @JG13_opcode: True enough - this is just a simple calculator for CLmax if Vmin is known - and visa versa. Or you can change the weight and see what happens. (This should be easy without the calculator since the formula is essentially CLmax=weight/X, so a 10% change in weight will create a 10% change in CLmax). So by implication it is the CLmax at Vmin: not much Mach effect at minimum stall speed. So it is only intended to be used as a quick way to check the sorts of calculations we have seen too much a lot of recently. Similarly, the calculation uses TAS as an input - if you want to see what the IAS for any specific plane or scenario is, you will have to enter the appropriate corrections yourself! The reason I did this is that it soon becomes apparent that with very slight changes in the assumptions for weight and Vmin, you can create a fairly wide range of outputs for CLmax. Hence the derived CLmax is only as good as the assumptions fed into the calculation. If these are derived from fairly approximate reports of IAS at stall speed and a guestimate for weight, for instance, you can generate "good agreement" with almost anything you want. Since one of the scenarios I checked was the calculations with Crumps assumptions he is entirely right of course, garbage in, garbage out. Calculating a CLmax using a fully laden aircraft and a TAS derived from an IAS rounded to the nearest 5mph will give you one answer - take out the weight of ammo and 50% of the fuel, model a 5mph error in the IAS - get a very different answer. The formula that the calculator uses is exactly the same as the one Crump has used in several posts, btw, just expressed in SI units.
Crump Posted October 15, 2016 Posted October 15, 2016 Similarly, the calculation uses TAS as an input Not the airspeed used in the formula I presented. That is why your results are not correct. The formula that the calculator uses is exactly the same as the one Crump has used in several posts, btw, just expressed in SI units. For example...that formula is specifically designed to use KNOTS in the BGS system. If you do not what is applicable and what is not....you will get a nonsensical result. I have not looked at your spreadsheet, either and I will not download anything from you either.
unreasonable Posted October 15, 2016 Author Posted October 15, 2016 (edited) I have not looked at your spreadsheet, either and I will not download anything from you either. Says it all really, yet you assert that it is incorrect. Even though the formula I use, when fed with your assumptions, returns your answer! Anyone is free to try that for themselves. As a reminder from post 143 in the locked thread: Crump said: Using the BGS system and a conversion factor for knots... EAS means sigma (density ratio) = 1 Velocity in KEAS = SQRT{295 (8580lbs) / (1.58 * sigma * 197 ft^2) Velocity equals = 90.2 Knots Equivilent Airspeed 90.2 KEAS * 1.15 = 104 mph EAS. 110 mph IAS - 104 mph EAS = Correction error of 6 mph. 104 mph EAS + .5 mph CEC = 104.5 mph CAS = 104.5 mph CAS + 5.5 mph PEC = 110 mph IAS as recorded by the RAE on WNr 313. First off, if you cannot see that that is just a different expression of the same physics then that is your problem - but then how could you know if you have not looked to see what it is? Anyone can input the following into my spreadsheet: Vmin 104 mph in kph = 167kph Calibration error 6mph in kph = 10kph - not necessary since Crump has already provided the correction, just for completeness. Weight 8,580lbs in kg = 3,892kg Density as it is by default, as this uses a value rather than being indexed as in Crumps formulation. Wing area = 18.3 m^2 Guess what the spreadsheet calculates? CLmax = 1.58! Finally, Crump, if you are not going to look at the attachment you have no business posting in this thread. Edited October 15, 2016 by unreasonable
Crump Posted October 15, 2016 Posted October 15, 2016 (edited) Says it all really, yet you assert that it is incorrect. Even though the formula I use, when fed with your assumptions, returns your answer! This is all my previous post says, If you do not what is applicable and what is not....you will get a nonsensical result. Anything else is something you have inferred. It is not intended. How can I say anything specific about something I have not examined? I cannot so why would you assume that is what I am doing? That is why I stated: I have not looked at your spreadsheet, either That formula does not take into account things like compressibility/Mach effects Listen to this..... The Clmax of 1.58 for the NACA 23015/23009 is the result of comparative studies set under known and specific conditions. That is sea level on a standard day at stall speed of the aircraft. If those conditions change and you do not properly account for those changes.....you will not get the same result and you will be modeling something else besides the intended aircraft. What that value does is set a known point in the mathematical system that is the aircraft and designed to be used with aircraft performance mathematics. One caveat for the reader....you are accounting for compressibility/mach effects in that performance math system by a general speed correction. That is your Compressibility Error Correction and your density effects thru the SMOE. That is a mathematical way to deliver the correct end results for aircraft performance purposes. The math is not designed to provide detailed analysis of flow dynamics, aircraft behavior, or airfoil theory. It is just ONE small part of an aircraft performance math system that tells you what the aircraft will DO. It is a solid foundation to examine what you should see within a reasonable margin of error (+ or - 3%) If the aircraft does not achieve that performance within a reasonable margin of error then it is flag to start looking for errors. First off, if you cannot see that that is just a different expression of the same physics then that is your problem - but then how could you know if you have not looked to see what it is? Ok unreasonable,I am going to assume you are really trying to seek the truth. It looks like you are holding velocity constant and therefore your CLmax must change. It is not. If our atmospheric conditions are steady at a constant altitude, then our CLmax will remain steady and velocity must change. That is the beauty of knowing the designers CLmax. We have fixed relationship and starting point under known conditions of flight. Edited October 15, 2016 by Crump
unreasonable Posted October 15, 2016 Author Posted October 15, 2016 (edited) The formula I am using is derived from http://aerostudents.com/files/introductionToAerospaceEngineering/introductionToAerospaceEngineeringFullVersion.pdf Please note 1.16 "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" This is a physical relationship expressed in physical quantities - no indexing required. It works neatly because it is expressed in Si units. For both of these reasons I prefer it to your formulation but they are mathematically equivalent. By simple algebra you can rearrange to solve for any of the variables if you know all of the others. This is what the calculator does. Air density is entered as a physical quantity at a standard day, or if you so choose, a different value for a test taken at altitude. Same with every other value - enter the appropriate physical value for whatever test you want to examine. The only tricky part is reconciling a reported IAS from a document with the TAS this formula needs. On the default corrections - I have just taken your figures. (Nearly all of the difference is in the PEC.) These can also be changed by the user if appropriate. " If our atmospheric conditions are steady at a constant altitude, then our CLmax will remain steady and velocity must change." Actually, If our atmospheric conditions and CLmax (and wing area and weight) all remain steady, the Vmin cannot change! How can it? It is a function of all of these things - unless you are saying that the formula is incorrect? The Vmin can only change if at least one other variable also changes. If you take the quantities for density, wing, weight and observed speed, and then make appropriate corrections to get the TAS, CLmax is mathematically derived. You are simply assuming that the "designers CLmax" is a given as a CLmax for the aeroplane in the test conditions - I am saying this assumption is not supported by the test data. The problem is that you can only get your 1.58 value from the RAE test data using a fully loaded weight and a possible low-ball estimate for the speed, both of which assumptions inflate the CLmax. AFAIK we simply do not know enough about the details of the test to be so categorical. It is possible that these were the values, but I find it highly unlikely that the weight assumption is correct, in which case either the Vmin was measured incorrectly and/or (corrected improperly by you), or the CLmax was not 1.58. The calculated CLmax is highly sensitive to both the weight and the speed data. This is why your repeated assertion that the RAE test gives good agreement with your wing CLmax is unconvincing. I can easily make it give a CLmax of 1.40 using entirely plausible weight assumptions alone! Lastly, we know from various wind tunnel tests that the measured CLmax of whole aeroplanes generally =/= CLmax of the wing section - it is somewhat less, depending on how closely an actual physical aeroplane resembles an idealized smooth model. Why should this case be an exception? Now for the last time, the idea of this thread is to give people a calculator to solve for the CLmax or Vmin, given inputs for weight, wing area and density, using the formula given above. They can then draw their own conclusions about how categorical it is advisable to be, given the sensitivity of the solutions to changes in the variables. Edited October 15, 2016 by unreasonable
Crump Posted October 15, 2016 Posted October 15, 2016 Actually, If our atmospheric conditions and CLmax (and wing area and weight) all remain steady, the Vmin cannot change! weight As weight changes...velocity will change unreasonable. You are simply assuming that the "designers CLmax" is a given as a CLmax for the aeroplane in the test conditions - I am saying this assumption is not supported by the test data. No assumption. It is a fact. It is not "test conditions", it is standard conditions. I can easily make it give a CLmax of 1.40 using entirely plausible weight assumptions alone! Because you are holding speed steady.... That is not how it works. Lastly, we know from various wind tunnel tests that the measured CLmax of whole aeroplanes generally =/= CLmax of the wing section - it is somewhat less, depending on how closely an actual physical aeroplane resembles an idealized smooth model. Why should this case be an exception? I am sorry but good agreement under known conditions is the foundation of aeronautical science. I did not go to college for some form of mysticism. It is a hard science. The airfoil section data agrees with the 3 dimensional wing agrees with the model agrees with the full scale model and agrees with the flying airplane given proper construction and finish. That is how airplanes are designed and work. There are things that will change that dynamic. There are called problems and the general result is they get solved. Sometimes it construction issues down at the factory, sometimes it is bugs getting into the wind tunnel intake and raising your surface roughness..... When things do not agree. It is a problem and you start looking for the answer. What i find funny is those people who have no formal training in aeronautical sciences but seem to think they know more about this than the folks who built the airplane! 1
unreasonable Posted October 15, 2016 Author Posted October 15, 2016 (edited) I am not "holding speed steady" I am using the reported number as you did, using all the other data we know - or guess - from the test, and then calculating the CLmax that must have held if the other data is correct. The only difference is that - unlike you - I have tested for sensitivities. The fact is that: a) You do not know the exact weight of the aircraft at the time the IAS was given. b) You almost certainly only know the IAS to the nearest 5mph, since it is highly unlikely that all of the reported speeds were precisely in multiples of 5mph. To make your numbers add up you have to assume that the weight was fully loaded, and that the IAS was not rounded down, possibly by nearly 5mph. If any/all of these are out, 1.58 is impossible. "General agreement" =/= "exact agreement"! Can you not even read your own sources? Just above the highlighted part of the last extract it says "The predicted maximum lift coefficient for the wing will be somewhat lower than the maximum lift coefficients of the sections used because of the non-uniformity of the spanwise distribution of lift coefficient. The difference amounts to about 4-7 percent for a regular wing with an aspect ratio of 6." So that is 4-7% in that example between sections and wing - never mind between wing and whole aeroplane. The difference between a CLmax of 1.58 and 1.42 is just 10%. Honestly Crump, just leave it alone and let people make up their own minds using whatever tools they want. The developers certainly will, and while I hope they take into account the new documentary evidence people have found for them, I doubt they will enjoy your lectures. Having you grandstanding in my thread is most unwelcome. Go away, or I will pester you in your "Community help needed" thread, which BTW seems awfully odd since unlike us gamers you are a professional pilot. I would have thought the kind of testing you are asking other people to do would be much better done by yourself. Actually tracks would be much better for your purpose, since you can look around and check instruments etc, change views, as you want rather than having to depend on where someone else choose to look. Edit - you really need not reply again. I know you always have to have the last word, but since this is my thread it would be common courtesy to make an exception in this case. Edited October 15, 2016 by unreasonable
Crump Posted October 15, 2016 Posted October 15, 2016 I will be happy to leave your thread. I just do not want you spouting aerodynamics principles that you do not know. It is not helpful to the community or the devs. Aerodynamics is much more than being able to work the lift formula, LOL. For example... Do you know what this is and what is used for?? a) You do not know the exact weight of the aircraft at the time the IAS was given. b) You almost certainly only know the IAS to the nearest 5mph, since it is highly unlikely that all of the reported speeds were precisely in multiples of 5mph. To correct the weight and airspeed data....
JG13_opcode Posted October 16, 2016 Posted October 16, 2016 Lift coefficients don't tell the whole story. Predicting the point of stall analytically is actually quite difficult, which is why wind tunnels and CFD are in widespread use. Cl varies with angle of attack and is affected by the deceleration of the aircraft.
unreasonable Posted October 16, 2016 Author Posted October 16, 2016 Lift coefficients don't tell the whole story. Predicting the point of stall analytically is actually quite difficult, which is why wind tunnels and CFD are in widespread use. Cl varies with angle of attack and is affected by the deceleration of the aircraft. My take on the formula I used was that like all scientific truths it is an approximation whose accuracy depends on the specifications of the conditions. My assumption was that it was a very good approximation at speeds close to Vmin, at least good enough so that any theoretical errors would be (much?) less than measurement errors, especially given the limited measurement accuracy. Is this reasonable? However I understand your point about the need to test - I had wondered why, if the CLmax was a given, anyone would bother with wind tunnels at all! Also, since CLmax is just a ratio and not a physical quantity, the best way of finding out what it is is to measure the actual physical quantities. Reification of indices, ratios etc being "A Bad Thing". But maybe I am just being old fashioned. On that note this formulation also forces the input of a specific quantity for air density: while you can start with the assumption of a "standard day", the result will only be accurate if your estimate is correct. So if you assume the flight test results were reported as observed at say 1,000m, (but with standard weather) you make a correction for altitude - which will give a higher CLmax value for any set of inputs than the standard. If you do not know the actual quantity at least you are forced to recognize this fact when estimating the reliability of any calculated results. I think the hardest part of using a model in this - or similar - form is the estimation of TAS. As I see it there are only two speeds that matter: the TAS which is the only speed that can have any actual physical consequences for the aeroplane, and a number reported in a document somewhere, which might be in a variety of different formats, and might contain measurement or even transcription error. You have to get from one to the other, and there is no general formula for how to do this AFAIK - each case has to be examined individually. On a side note, the more I look into these issues the more sympathetic I am for the developers who have to come up with one single model that "sort of agrees" with everyone's take on the data!
Crump Posted October 16, 2016 Posted October 16, 2016 My assumption was that it was a very good approximation at speeds close to Vmin, at least good enough so that any theoretical errors would be (much?) less than measurement errors, especially given the limited measurement accuracy. Is this reasonable? That is a very good observation and perfectly reasonable. It is also exactly why having a measured CLmax derived from comparative studies by the engineering team and the one used in the design is pure gold in terms of aerodynamic information for BoS. Spanwise flow will lower Clmax from the 2D data at high Reynolds Numbers (turbulent flow) but this is always determined experimentally in a variable density wind tunnel. This reduction for a rectangular wing(small taper but close enough for the FW-190) with an Aspect Ratio of 6 (6.08 = FW190) will fall somewhere between a 4% to a 7% reduction in CLmax from the measured airfoil data. Airfoils must be compared at the same Reynolds Number, so at 6.1 X 10^-6 Re which is the Re in the vicinity of the FW-190 wings stall speed, we can see the 2D data airfoil gives us a CLmax of 1.7. IIRC, 6.1 x 10^-6 is on a standard day represents turbulent flow. If we assume a worst case at the lower end our loss scale, then 7% reduction in the airfoils CLmax to reach the wings CLmax: 1.7 - (1.7 * .07) = 1.7 - .119 = 1.581 for the clean wing CLmax. Both Grumman and Focke Wulf wing designs fall exactly where all the data in the aeronautical sciences world for that airfoil selection says it should. 1
unreasonable Posted October 16, 2016 Author Posted October 16, 2016 I thought you had pushed off? I agree that 1.7 * 0.93 = 1.581 at least! However, since you have agreed that CLmax for model wing section =/= CLmax for wing, having kindly provided the authoritative source, you have already conceded part of my main point: a) CLmax for wing section =/= for wing =/= for smooth wooden plane scale model =/= for polished and sealed test aeroplane =/= for aircraft in operational conditions. b) The CLmax for the wing section represents the absolute maximum possible CLmax for the whole aeroplane. c) How much each of these difference amounts to will vary case by case, both by type and individual plane. d) The actual CLmax for any specific aeroplane (or type of aeroplane) needs to be validated experimentally. e) To get the RAE numbers to agree with a full CLmax of 1.58 is possible but requires an extreme set of assumptions. More reasonable assumptions are consistent with lower values. g) And finally! Having a handy calculator may help people to make up their own minds, which was the point of this thread, Crump, not to have this stimulating discussion with you. Finally I would note that the Reynalds number should be about 6,100,000 ie 6.1 * 10^6 not 10^-6 but I am just the "philosphy major" so what do I know.
Crump Posted October 16, 2016 Posted October 16, 2016 However, since you have agreed that CLmax for model wing section =/= CLmax for wing, having kindly provided the authoritative source, you have already conceded part of my main point: No I did not concede anything other that what the science says.... I simply pointed out the what the 2D airfoil data says our wing should be at and that everything after that agrees with it. The airfoil section data agrees with the 3 dimensional wing agrees with the model agrees with the full scale model and agrees with the flying airplane given proper construction and finish. Is just as true today as it was yesterday. To get the RAE numbers to agree with a full CLmax of 1.58 is possible but requires an extreme set of assumptions. More reasonable assumptions are consistent with lower values. You mean like the actual Position Error Curve and Compressibility correction. I would not call those "extreme assumptions"! They are actually kind of a requirement. http://www.spitfireperformance.com/bf274.html The actual CLmax for any specific aeroplane (or type of aeroplane) needs to be validated experimentally. Which is exactly what happened...LOL. Both Focke Wulf and Grumman did a very detailed analysis using the best scientific equipment and technology of the day. So your's is better and trumps their conclusions? da hierfür bereits vergleichsuntersuchungen der FW190 vorhanden wraen , since comparative studies of the FW190 already exist. "Comparative Studies" has a very specific meaning in engineering. It is where we compare differing methodology such as wind tunnel data for models and full scale aircraft, airfoil data, and all available resources. It is the value that all methods of measuring agree upon! The results from the two comparative studies showed that there was in general good agreement between CFD and wind-tunnel measurements. http://papers.sae.org/2014-01-2443/ This is the result of comparative studies on the FW-190 wing design conducted by Focke Wulf. That is comparing different MEASUREMENT METHODS. If they agree, then the data is valid. If they do not, then back to the drawing board. Once more, a modern CFD VSAERO analysis returns the same result! Do you know how VSAERO works? You do not input the airfoil data or calculate coefficients of lift. You build a computer model of the design and VSAERO uses Computational Fluid Dynamics to analyze it. You can set whatever surface roughness and other parameters you want. VSAERO couples integral methods of potential and boundary layer flows for low runtimes – full configurations in as little as five minutes. Compatible with structured, unstructured and hybrid surface meshes, VSAERO can also be expanded to models with 50,000 or more surface panels for extremely complex models. Besides standard surface calculations, flowfield properties are computed for off-body velocity surveys and on/off-body streamlines. The ability to calculate internal and external flows, non-uniform inflow and body rotation, makes VSAERO applicable to fluid flow problems in the aerospace, automotive and marine industries. Extensive options for wake modeling, surface modeling and matrix solving allow the user to fine tune the code to best match their needs. Special purpose modules FSWAVE and ROTOR expand VSAERO’s simulation capabilities to include nonlinear hydrodynamic wave effects on ships and helicopter rotor/ fuselage interactions. Zonal coupling to Navier-Stokes codes is also available. Running on a wide variety of platforms, including desktop PCs, VSAERO is used worldwide. VSAERO has been used in the development of Rutan Voyager and Beech Starship aircraft, the Stars and Stripes racing yachts and the Sunraycer solar automobile. It is the pure gold in terms of aerodynamic data and is the ONLY calculation methodology that is considered as good as actual measurement. And finally! Having a handy calculator may help people to make up their own minds, which was the point of this thread, Crump, not to have this stimulating discussion with you. I think it is fantastic that you are trying to learn this. Just keep it in perspective and do not think you can overturn the designers conclusion just because you built an openoffice spreadsheet to solve the lift formula. 1
3./JG15_Kampf Posted October 16, 2016 Posted October 16, 2016 (edited) I think this discussion of the CLmax will only end when someone catch an original fw190 (still exist?) And put it in a modern wind tunnel and analyze. I think the results surprise many. If it is possible I do not know. LOL Edited October 16, 2016 by JAGER_Kampf
Crump Posted October 16, 2016 Posted October 16, 2016 I think this discussion of the CLmax will only end when someone catch an original fw190 (still exist?) And put it in a modern wind tunnel and analyze. LOL...maybe but I am willing to bet they will get the same results as Focke Wulf, GmbH and VSAERO!! Cause that is exactly what they did!
Hutzlipuh Posted October 16, 2016 Posted October 16, 2016 (edited) I think this discussion of the CLmax will only end when someone catch an original fw190 (still exist?) And put it in a modern wind tunnel and analyze. I think the results surprise many. If it is possible I do not know. LOL somebody should contact flugwerk...they were building a 190 after original plans...just with a ash 82 engine :D http://www.flugwerk.de/html/page.php?GID=19&SID=4 Edited October 16, 2016 by Hutzlipuh
JG13_opcode Posted October 16, 2016 Posted October 16, 2016 (edited) Also, since CLmax is just a ratio and not a physical quantity, the best way of finding out what it is is to measure the actual physical quantities. Reification of indices, ratios etc being "A Bad Thing".Gotta disagree on this point. Coefficients are an accepted and widely used practice in engineering. Not just aerospace, either. The fact that something is a ratio does not make it inherently less accurate or valid. Edited October 16, 2016 by JG13_opcode
Crump Posted October 16, 2016 Posted October 16, 2016 Focke Wulf already did all of this quite extensively... By December 1944, they might have an idea what the CLmax of their design is..... They might have figure it out by then if they were some 16% off on their measurements. 1
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 Gotta disagree on this point. Coefficients are an accepted and widely used practice in engineering. Not just aerospace, either. The fact that something is a ratio does not make it inherently less accurate or valid. They are used in finance and economics too - no denying that they are useful, especially in heuristics where you want usable approximations in a hurry. From the point of view of someone learning a topic they can be a problem because they distract attention from the underlying measurable quantities. They can be a problem if someone starts to assume that they have an independent existence - it is not that they are less accurate and valid, but that their accuracy and validity is a function of the underlying physical quantities, not the other way around. What Crump is saying is that in the RAE test the weight and Vmin MUST be certain quantities because he KNOWS the CLmax of this specific aeroplane. I say this is untrue and bad science. It might be an acceptable engineering heuristic but it would get you laughed out of any science lab. From my own point of view the purpose of this thread was to see if there were any mistakes in the calculator or possible improvements. No-one has come up with any mistakes in the formulas - so far. There have been a couple of questions by PM about data entry that suggested that improving the clarity of what was required for each field might help so I will look at that, thank you for constructive feedback. Given that Crump has reappeared with his endless spam of stuff we have all seen before, despite saying that he would go away, it is time to end this thread. MODERATORS PLEASE LOCK
=EXPEND=13SchwarzeHand Posted October 17, 2016 Posted October 17, 2016 I´m not sure if it is a viable way to have threads closed, just if the discussion isn´t going the way the OP wants it. Sure I can see how people are annoyed by some posters, but that's part of posting on a forum. Crump is a member of this forum, so as long as he is not banned, he may post his stuff if he pleases. Threads should not become the "property of the OP". They are dynamic processes and you have to live/deal with all sorts of posts. That's part of the game. I think the forum is going the wrong way if we make this the usual procedure. 6
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 I´m not sure if it is a viable way to have threads closed, just if the discussion isn´t going the way the OP wants it. Sure I can see how people are annoyed by some posters, but that's part of posting on a forum. Crump is a member of this forum, so as long as he is not banned, he may post his stuff if he pleases. Threads should not become the "property of the OP". They are dynamic processes and you have to live/deal with all sorts of posts. That's part of the game. I think the forum is going the wrong way if we make this the usual procedure. The problem with this is that whenever someone posts anything that Crump believes might possibly be contrary to his views we get the same reams of posts that we have all seen before, delivered with his usual combination of condescension and insult. The topic of the thread was about the calculator - which Crump refused to look at but asserted was wrong - even though it returns the same results from the exactly the same inputs as his own formula. At no point has he addressed the topic - the calculator, the usefulness of the formula or the sensitivities inherent in the calculations. Assuming that he is truthful, he still has not even looked at it. Or perhaps he was lying when he said he would never download anything I posted. Either way this is hardly conducive to constructive discussion. All he has done is simply to reassert, endlessly, that he knows the CLmax is 1.58 For all know perhaps it is, but that is not the point - the point is the sensitivity of the calculated outputs to the assumptions used. My mistake perhaps for replying to him on the specifics of the Fw190 case - but it is difficult to ignore someone whose clear intention from the start is to trash the thread. If you think it is reasonable for someone to a) say they will never look at the material the thread is about, and b) claim that it is wrong, then perhaps this is all part of the give and take of communication for the modern internet generation. God help us. 1
Mastermariner Posted October 17, 2016 Posted October 17, 2016 I think this discussion of the CLmax will only end when someone catch an original fw190 (still exist?) And put it in a modern wind tunnel and analyze. I think the results surprise many. If it is possible I do not know. LOL I know where we can get some, just bring a shovel! http://alert5.com/2016/10/17/50-fw190-a-3s-buried-in-turkey/ Master
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 I read about that too - perhaps we should have a whip round and buy one! On the subject of calculators, as I said the CLmax calculator does not correct from some reported speed to the TAS which the formula uses - you have to do that yourself. There is a calculator here that does that, and also gives the standardized atmosphere density changes. What it does not do as far as I can see is incorporate PEC, presumably because that is specific to each aeroplane design. I am not giving a warranty that it is correct, but it might be useful to someone. https://sites.google.com/site/maltapplication/home
Crump Posted October 17, 2016 Posted October 17, 2016 which Crump refused to look at but asserted was wrong Actually I did look at. Half of it did not work.
Crump Posted October 17, 2016 Posted October 17, 2016 What it does not do as far as I can see is incorporate PEC, presumably because that is specific to each aeroplane design. They are specific to each installation. While there are general trends in each pitot static system and the more data points one has, the more accurate the PEC curve. The same design pitot static system will have a similar trend on the same type aircraft. That is how a general PEC curve is established. The most accurate is constructing a specific PEC curve for a specific airplane. Pitot static systems have to be maintained. Leaks in the system will cause inaccuracies beyond the normal PEC curve. Leaks are common. The airframe flexs and normal expansion and contraction will cause fittings to leak. That is why there is a requirement to check the Pitot Static system for leakage on a periodic basis for Instrument flight rules rated aircraft. Crump believes might possibly be contrary to his views It is not "my view", this is hard science not extensionalism, "My view" is the manufacturer was much more competent than you or me when it comes to the engineering details of their design. They had the tools, knowledge, and experience to measure. Since it is not World War I but rather a time when the engineers were very well versed in subsonic aircraft aerodynamics and their conclusions fit nicely into everything we know today, there is no reason not to use it as representative of a properly finished and constructed aircraft. It is a simply fact that a Clmax of 1.58 is representative of the clean wing stall AoA.
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 Actually I did look at. Half of it did not work. Stop fibbing. My copy works fine. (Anyone else have a problem?) Enter any set of input data in the top half and solve for CLmax. Enter the same density, weight and wing data in the bottom half, using the CLmax from the top half as an entry and solve for Vmin. It will return the same Vmin as you entered in the top half - it simply the same formula. It really is not complicated. You could add sections solving for weight, wing area and air density if you want - it is just algebra.
Crump Posted October 17, 2016 Posted October 17, 2016 (edited) Stop fibbing. My copy works fine. (Anyone else have a problem?) Mine does not. I do not have open office and there maybe an issue with your open office importing to Office. That being said, much more complicated spreadsheets have imported just fine. I have not said your spreadsheet was wrong. The lift formula is very easy and is generally discussed the third topic of the course on day 3 of Aerodynamics 101, after talking about units, the atmosphere, and airspeed measurement. Edited October 17, 2016 by Crump
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 If it opens it should all work - there is no condition I can think of that would lead to half of it not working. If it will not really open in your version of MS Office there is nothing I can do about that as I do not have MS office installed. Other people have opened it OK. Any updates on how discussions with devs are going? I have no idea - I am not a dev-pesterer. You may be better off asking in one of the other threads, probably anyone who knows the answer to that has given up on this thread long ago....
-TBC-AeroAce Posted October 17, 2016 Posted October 17, 2016 (edited) You are both using two different theorys to calculate CLmax. One is via Prantal's lifting line theory which uses the geometric shape of a 2D aerofoil to caluclate 2D aerodynamic forces and then it uses the wing geometry to extrapilate the 2D forces to a 3D model. This is 100% theory based on circulation. The other method is using Bernoulli's equation which is more of an experimental calculation based on real life numbers. Both methods are used IRL, Prantal more for the initial design of a aerofoil/wing and Bernoulli's as more of a later performance check. Although both are used in the design process in a kind of feed back loop manner until there is good convergence. Basically Prantal will tend to over estimate CLmax as it is not taking into account a lot of things(although it can to some extent) such as fuelalarge drag and much more (It only cares for the wing). Bernoulli's will give a more correct answer if it's inputs are correct. In fact Bernoulli's in theory cant be wrong if the inputs are correct. PS Compressible effects only start to happen at a Mach number of 0.3 and only really have a big impact approaching the trans sonic range. Air below Mach 0.3 is considered incompressible i.e the stalling speed of a ww2 fighter should not have any real compressed flow. PPS alot of this back and forward might be because one of you is thinking about the CL of the aircraft and one is thinking about the CL of the plane! They are of course very different Edited October 17, 2016 by AeroAce 1
Crump Posted October 17, 2016 Posted October 17, 2016 Unreasonable, You seem to think I do not want you to be able to work the math. On the contrary, I am happy that you can do it and hope that others will also follow suit. It will give some insight into the complexities being discussed and maybe even cut down on the "white noise" of semi or ill-informed posters. That is of course as long as people realize there is a lot more to this than just working the lift formula and simply use this as an exercise to illustrate just how much they do not know instead of deciding they can now overturn the results of engineering design teams.
Crump Posted October 17, 2016 Posted October 17, 2016 You are both using two different theorys to calculate CLmax. One is via Prantal's lifting line theory which uses the geometric shape of a 2D aerofoil to caluclate 2D aerodynamic forces and then it uses the wing geometry to extrapilate the 2D forces to a 3D model. This is 100% theory based The other method is using Bernoulli's equation which is more of an experimental calculation based on real life numbers. Both methods are used IRL, Prantal more for the initial design of a wing and Bernoulli's as more of a later performance check. Although both are used in the design process in a kind of feed back loop. Basically Prantal will tend to over estimate CLmax as it is not taking into account a lot of things(although it can to some extent). Bernoulli's will give a more correct answer if is's inputs are correct. In fact Bernoulli's in theory can be wrong if the inputs are correct It is not even close to being that far Aero.... They are just working the basic Lift Formula. There is not a single method of calculation that is more accurate or even considered in the face of measured results or CFD analysis. The documents I posted are from measured data and CFD analysis. What I am using Newtonian based aircraft performance math that explores the edges of the performance envelope. Lifting Line theory is behavior math. It will tell you how the aircraft behaves but nothing about the actual performance envelope. That has to be measured in lifting line theory and the model adjusted to conform with those measurements. That CLmax measured by Focke Wulf purpose was to be used in Newtonian based performance math.
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 You are both using two different theorys to calculate CLmax. One is via Prantal's lifting line theory which uses the geometric shape of a 2D aerofoil to caluclate 2D aerodynamic forces and then it uses the wing geometry to extrapilate the 2D forces to a 3D model. This is 100% theory based The other method is using Bernoulli's equation which is more of an experimental calculation based on real life numbers. Both methods are used IRL, Prantal more for the initial design of a aerofoil/wing and Bernoulli's as more of a later performance check. Although both are used in the design process in a kind of feed back loop. Basically Prantal will tend to over estimate CLmax as it is not taking into account a lot of things(although it can to some extent) such as fuelalarge drag and much more (It only cares for the wing). Bernoulli's will give a more correct answer if it's inputs are correct. In fact Bernoulli's in theory cant be wrong if the inputs are correct. PS Compressible effects only start to happen at a Mach number of 0.3 and only really have a big impact approaching the trans sonic range. Air below Mach 0.3 is considered incompressible In Crumps earlier posts he was using a version of the same equation to derive a Vmin and weight (+ imputed rho + S) that was consistent with his assumption that the CLmax was 1.58 My calculator was merely a computational aid to test the sensitivities of this equation - I have no idea if the equation is even right, but I take it on trust from my quoted source as being OK at this sort of speed. I am certainly not claiming to be an aeronautics expert! But I can do algebra and check the data. Also I know the difference between an assumption and an empirical fact. Your point that the lifting line theory may tend to overstate the result is certainly consistent with some of the data that has been presented in various threads, so I certainly find that plausible.
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 Unreasonable, You seem to think I do not want you to be able to work the math. Where on earth did you get that idea? I do not give a monkey's whether you want me to work the maths or not. First you say you were not going to open any attachment from me - if you do not think that is an insult then you must be particularly thick skinned. Then you say you have opened it but "half of it does not work". Tell the marines.
Crump Posted October 17, 2016 Posted October 17, 2016 There is no correlation required between 2D data and lifting line wing CLmax. The math is not designed to predict PERFORMANCE but rather BEHAVIOR. Lifting line theory gives good agreement for in flight behaviors for high aspect ratio wings in the subsonic realm. The stall and post stall regime however is totally empirical in lifting line. All the edges of the envelope are empirical.
-TBC-AeroAce Posted October 17, 2016 Posted October 17, 2016 (edited) I guess my years of studying Aerodynamics to masters level was a waste then. I will look at it in detail instead of off the top of my head and check back. PS I would be very grateful if someone had the correct wing/aerofoil geometry for the 190-A series so I can run some numbers. I can even do a CFD run if more detailed geometry is at hand Edited October 17, 2016 by AeroAce
unreasonable Posted October 17, 2016 Author Posted October 17, 2016 Lifting line theory and behavior math is then adjusted to MEET that newtonian based measured CLmax. NOT THE OTHER WAY AROUND!!! What you did, Crump, in the discussion of the RAE results, was to derive a Vmin based on an assumed CLmax of 1.58. Given a fixed wing area and standard rho, you found the only possible points for W and Vmin that could work (given that W cannot be over fully loaded weight), to be consistent with 1.58 Nothing algebraically wrong with that, assuming these inputs, but it is not empirical discovery of the CLmax. Here is the relevant extract from your post again to refresh your memory. Using the BGS system and a conversion factor for knots... EAS means sigma (density ratio) = 1 Velocity in KEAS = SQRT{295 (8580lbs) / (1.58 * sigma * 197 ft^2) Velocity equals = 90.2 Knots Equivilent Airspeed 90.2 KEAS * 1.15 = 104 mph EAS. 110 mph IAS - 104 mph EAS = Correction error of 6 mph. 104 mph EAS + .5 mph CEC = 104.5 mph CAS = 104.5 mph CAS + 5.5 mph PEC = 110 mph IAS as recorded by the RAE on WNr 313. You did not ask yourself what the CLmax would have been, given a reasonable range of estimates for the inputs, which really would have been empirically measuring the CLmax. Had you done so realistically you would have found that there is considerable uncertainty over both the observed speed and weight, enough to give a fairly wide range of answers. (Perhaps not 1.17).
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