Aces High Bulletin Board
General Forums => Aircraft and Vehicles => Topic started by: gripen on September 09, 2004, 04:26:00 AM
-
In the "Draining E in turns" thread rose a question if the Oswald's efficiency factor (know also as e factor) can be estimated from the unknown wing by knowing only the aspect ratio of that wing. In other words, is it possible to estimate the lift distribution of the wing by knowing only the aspect ratio of it?
So I open a little contest: The first one who can explain if this is possible or not with sensible logic will receive a pdf copy of the RAE report called: "Notes on the Dog Fight" from me. As a bonus prize the winner will also get a flight tested German report on the Bf 109G/AS for performance analysis.
gripen
-
I'd say NO.
Aspect ratio is only a function of chord and span.
You'd be missing shape, wingtip shape, thickness and dihedral.
What's an E factor anyway :D
-
I dont think you can.
proof:
use 2 wings,
A got a profile 1cm thik 400cm deep, square.
B one got a profile 400cm thik, 1cm deep
now with same aspec ratio these 2 wings have diffrent factors.
You say second one is no wing? well but then you define a profile for what is a wing and what not.
So that is the proof you need to know more than the aspect ratio.
-
Here's the problem - the efficiency of a wing depends on it's planform and crosssection.
The aspect ratio tells you something about the planform, but it is not even a sufficient statistic for that-compare a straight taper vs. an elliptical wing for example.
Applied physicists figured out most of the effects of planform on the efficiency of wings very early (say by the 1920s at least). It took many decades to work out the effects of different cross-sections and even today some of the results are approximations.
That is why NACA spent two decades conducting wind tunnel tests of airfoil cross sections and publishing these results. The principal was called parameter variation-you don't know the exact formula but you try to approximate it by gathering data generated by small changes in the airfoil.
Consider the P51 wing with and without a laminar flow cross section. The planform does not change, so the aspect ratio is the same. But, in principal at least, the laminar cross section is more efficient.
-Blogs
-
Joeblogs winns the cooky! :)
-
Here is what a guy that teaches aeronuatics told him:
------------------------------------------------------------------------------------
quote:
--------------------------------------------------------------------------------
Originally posted by gripen
This is the only one I have in the hand of those you mentioned and the graphs in the page 73 actually give correction factors of the rectangular wing against aspect ratio.
--------------------------------------------------------------------------------
Correct, Figure 2-44 (b) shows the relationship of those factors to Aspect Ratio for a rectangular wing, and that illustrates the dependency of e to Aspect Ratio. In the previous graph you see the influence that taper ratio has for an aspect ratio of 6. There are other formulae that also include the taper ratio so that values for non rectangular wings can be estimated also.
quote:
--------------------------------------------------------------------------------
In the previous page the writers actually say that:
"For an untwisted wing with elliptical planform shape these correction factors are zero"
--------------------------------------------------------------------------------
Correct, and that’s why the elliptical planform theoretically has the least induced drag, in that case the efficiency factor e would equal one.
quote:
--------------------------------------------------------------------------------
If I understand this correctly those formulas you gave above work only for the rectangular wings and we can't determine the shape of the wing from the aspect ratio only.
--------------------------------------------------------------------------------
Partly correct… It is true that formulae that only include aspect ratio only apply to a single taper ratio, it just happens that the graph in Perkins & Hage is for zero taper. The previous graph on the same page illustrates the relationship for both taper ratio and aspect ratio but even so the dependency upon aspect ratio is clear.
However, the formulae presented by Professor Wood is for a taper ratio of 0.57 which yields an almost elliptical lift distribution, which is why it fits the experimental data so closely over a wide range of examples. Most WWII aircraft had a taper ratio close to that value (Me109 was approx 0.52) because the designers knew about the benefits of elliptical lift distribution and that it could be achieved quite closely with a wing of that taper, that’s why a formulae that only includes aspect ratio could still be of such good practical use. However, other graphs and formulae that include both taper ratio and aspect ratio were common, today it is just as easy to include everything, including twist, camber, and sweep etc.
Hope that helps
Badboy
------------------------------------------------------------------------------------
Crumpp
-
I did not actually see the discussion of this e variable, but the wider discussion sure sounded to me like it is an application of the mathematics for lift and drag most commonly characterized in terms of the "wetted" portion of a flat plate exposed to the slip stream.
If that is an accurate characterization of the discussion, this is all about measuring the properties of the planform. The effects of cross-section (and other things) are typically subsumed in a constant. In other words, the effect of the shape of the cross section has to be explained by an additional set of equations.
-Blogs
Originally posted by Crumpp
Here is what a guy that teaches aeronuatics told him:
------------------------------------------------------------------------------------
quote:
--------------------------------------------------------------------------------
Originally posted by gripen
This is the only one I have in the hand of those you mentioned and the graphs in the page 73 actually give correction factors of the rectangular wing against aspect ratio.
--------------------------------------------------------------------------------
Correct, Figure 2-44 (b) shows the relationship of those factors to Aspect Ratio for a rectangular wing, and that illustrates the dependency of e to Aspect Ratio. In the previous graph you see the influence that taper ratio has for an aspect ratio of 6. There are other formulae that also include the taper ratio so that values for non rectangular wings can be estimated also.
quote:
--------------------------------------------------------------------------------
In the previous page the writers actually say that:
"For an untwisted wing with elliptical planform shape these correction factors are zero"
--------------------------------------------------------------------------------
Correct, and that’s why the elliptical planform theoretically has the least induced drag, in that case the efficiency factor e would equal one.
quote:
--------------------------------------------------------------------------------
If I understand this correctly those formulas you gave above work only for the rectangular wings and we can't determine the shape of the wing from the aspect ratio only.
--------------------------------------------------------------------------------
Partly correct… It is true that formulae that only include aspect ratio only apply to a single taper ratio, it just happens that the graph in Perkins & Hage is for zero taper. The previous graph on the same page illustrates the relationship for both taper ratio and aspect ratio but even so the dependency upon aspect ratio is clear.
However, the formulae presented by Professor Wood is for a taper ratio of 0.57 which yields an almost elliptical lift distribution, which is why it fits the experimental data so closely over a wide range of examples. Most WWII aircraft had a taper ratio close to that value (Me109 was approx 0.52) because the designers knew about the benefits of elliptical lift distribution and that it could be achieved quite closely with a wing of that taper, that’s why a formulae that only includes aspect ratio could still be of such good practical use. However, other graphs and formulae that include both taper ratio and aspect ratio were common, today it is just as easy to include everything, including twist, camber, and sweep etc.
Hope that helps
Badboy
------------------------------------------------------------------------------------
Crumpp
-
I did not actually see the discussion of this e variable, but the wider discussion sure sounded to me like it is an application of the mathematics for lift and drag most commonly characterized in terms of the "wetted" portion of a flat plate exposed to the slip stream.
You are correct. Badboy simply points out that Oswalds Efficiency factor is a much more accurate way of determining the efficiency of a wingtip than just estimating it.
Gripen contended that any formula (Oswalds) that includes AR has nothing to do with the wingtip efficiency as he was estimating the value.
The discussion came up after this article was presented:
http://www.thetongsweb.net/AH/EAAjanuary1999.pdf
Using Oswalds Formula, the FW-190 and the Spitfire are very close in their wingtip efficiency with the Spitfire having a slightly better value as is confirmed by this article. It is not nearly as large as many of the estimates. Reason being the physical shape of the Spitfire's wingtip is not quite the same as the wet area lifting surface. In Short, the more detailed your calculations the more Lednicers conclusions are confirmed.
This leads to the FW-190 having less drag than the Spitfire in many portions of the flight envelope. The Spitfire gains advantage in low speed manuvering and pulling hard turns.
In level flight the faster you go the more the 190 moves to advantage, in the zoom climb, and in turning below around 4 G's and high speed.
You can see the whole the thread here:
http://www.hitechcreations.com/forums/showthread.php?s=&threadid=128143&perpage=50&pagenumber=5
Great discussion with lots of good information in it.
Crumpp
-
aspect ratio is not the whole story. consider a triangular wing, swept wing and rectangular wing.
The "e" factor is where all the errors of the models are burried. It is a fuzziness factor that can be estimated in the design stage but better be tested when the wing is actually built.
Bozon
-
Definately agree Bozon.
Crumpp
-
An interesting application of the e factor is found at
http://www.aerospaceweb.org/question/aerodynamics/q0184.shtml
I was wrong - Oswald's e takes into account efficiency losses that result from departures from a perfect elliptical planform.
The equation for total drag contains two unknowns. The first is a proportionality constant and the second is a "minimum 3d drag coefficient." What's interesting about the latter is that the author resorts to numbers derived from actual wings. In other words, he needs data (or an equation) that is outside of the lifting line equation he is working with.
-Blogs
Originally posted by joeblogs
Here's the problem - the efficiency of a wing depends on it's planform and crosssection.
The aspect ratio tells you something about the planform, but it is not even a sufficient statistic for that-compare a straight taper vs. an elliptical wing for example.
Applied physicists figured out most of the effects of planform on the efficiency of wings very early (say by the 1920s at least). It took many decades to work out the effects of different cross-sections and even today some of the results are approximations.
That is why NACA spent two decades conducting wind tunnel tests of airfoil cross sections and publishing these results. The principal was called parameter variation-you don't know the exact formula but you try to approximate it by gathering data generated by small changes in the airfoil.
Consider the P51 wing with and without a laminar flow cross section. The planform does not change, so the aspect ratio is the same. But, in principal at least, the laminar cross section is more efficient.
-Blogs
-
If you read Oswalds paper where he introduces the "e" factor, he was deriving it from flight tests and it is supposed to consider the efficiency of the whole aircraft, not just the wing. Link: http://naca.larc.nasa.gov/reports/1933/naca-report-408/index.cgi?thumbnail1#start
-
That's true.
The efficiency factor in that article appears in a formula that tries to correct for calculations of parasitic drag that assume an elliptical planform. When the actual wing does not take that shape, an adjustment is made, which essentially reallocates some drag from parasitic to induced drag. The efficiency factor is introduced in that calculation.
Getting back to Grippen's point, though, Oswald's e is used to correct for departures from eliptical planform and is therefore not just a function of the aspect ratio.
In Oswald's paper, he's backing actual numbers out from data on aircraft weight, span loading, and span. This is the equivalent of calculating an ideal for a given amount of power (controlling for propeller efficiency) and calculating the distance from the ideal from flight test data.
-Blogs
Originally posted by justin_g
If you read Oswalds paper where he introduces the "e" factor, he was deriving it from flight tests and it is supposed to consider the efficiency of the whole aircraft, not just the wing. Link: http://naca.larc.nasa.gov/reports/1933/naca-report-408/index.cgi?thumbnail1#start
-
Getting back to Grippen's point, though, Oswald's e is used to correct for departures from eliptical planform and is therefore not just a function of the aspect ratio.
Gripen is not correct. An instructor that teachs aeronautics explained it to him.
Read the third paragraph down and it explains Oswald's e factor and it's use.
http://naca.larc.nasa.gov/reports/1933/naca-report-408/index.cgi?page0001.gif
A correction is made for the variation of parasitic resistance with angle of attack and nonelliptical wingloading by including in the induced drag term a factor e, called the "airplane efficiency factor". The correction is thus assumed to be proportional to CL^2.
Now check out the induced drag formula.
As Badboy explained:
Correct, Figure 2-44 (b) shows the relationship of those factors to Aspect Ratio for a rectangular wing, and that illustrates the dependency of e to Aspect Ratio. In the previous graph you see the influence that taper ratio has for an aspect ratio of 6. There are other formulae that also include the taper ratio so that values for non rectangular wings can be estimated also.
And for using AR as part of the formula to solve for "e":
Most WWII aircraft had a taper ratio close to that value (Me109 was approx 0.52) because the designers knew about the benefits of elliptical lift distribution and that it could be achieved quite closely with a wing of that taper, that’s why a formulae that only includes aspect ratio could still be of such good practical use.
Crumpp
-
Well, if we have no idea about taper ratio, there is no way we can estimate the e factor with the aspect ratio only. However Badboy's poin't is quite sensible; most rectangular winged WWII fighters had quite similar taper ratio so in many cases generalized formula do quite well.
I have calculated some e factor data from the drag polars I have. These are e factors for whole airplane as Oswald and Perkin&Hage handle it:
Bf 109G, this is from the VL (Finnish State aircraft factory) drag polar:
Cl 0,1 => 0,741658161
Cl 0,2 => 0,798708788
Cl 0,4 => 0,769126981
Cl 0,6 => 0,759747384
Cl 0,8 => 0,787352739
Cl 1,0 => 0,769126981
Cl 1,2 => 0,713350597
Cl 1,4 => 0,619327448
The average is about 0,75
I have also a polar from Messerschmitt AG (I had to approximate Cd0):
Cl 0,1 => 1,038321425 (not in average)
Cl 0,2 => 0,83065714
Cl 0,3 => 0,778741069
Cl 0,4 => 0,755142854
Cl 0,5 => 0,70156853
Cl 0,6 => 0,692214283
Cl 0,7 => 0,696955477
Cl 0,8 => 0,692214283
Cl 0,9 => 0,7067566
Cl 1,0 => 0,711179058
The average 0,73
Mustang, RAE wind tunnel data:
Cl 0,6 => 0,789174119
Cl 0,4 => 0,814631349
Cl 0,2 => 0,717431018
The average is about 0,77
I also tried to calculate from the NACA 916 but there is so much varition that sensible analysis seems to be quite impossible; the e factors varied from 0,5 to 1,3 or something.
Regarding prizes; well I think the answer is not possible but in certain conditions it might be possible as Badboy pointed out.
So I think I'll send the prize to Angus and Blogs and Schutt will receive the Bonus, so post your adress and I'll send the docs. If Badboy is listening I can send both documents to you too.
gripen
-
If you read through all of the derivations, you see e is defined as a correction to parasitic drag (which is independent of the coefficient of lift while induced drag is not) that results when the planform is not ellipitcal - hence the reference to Munk's span adjustment. In all the subsequent derivations, span is adjusted by a factor (k) that reflects this.
At the end of the day the effect of this is a change in a scaler in an equation that is, among other things, a function of span. Hence the function is not soley deterimined by the aspect ratio.
Yes the function can be inverted to define the aspect ratio as a function of e, but that wouldn't mean that aspect ratio is a function of e alone.
The fact that one must assume a particular planform-rectangular with a specific taper-in order to draw a graph of the function makes this clear. One could assume different planforms and draw different graphs. The graphs would look different even if the aspect ratio was the same.
Grippen asked is aspect ratio a sufficient statistic to calculate Oswald's e? The answer is clearly no.
-blogs
Originally posted by Crumpp
Gripen is not correct. An instructor that teachs aeronautics explained it to him.
Read the third paragraph down and it explains Oswald's e factor and it's use.
http://naca.larc.nasa.gov/reports/1933/naca-report-408/index.cgi?page0001.gif
Now check out the induced drag formula.
As Badboy explained:
And for using AR as part of the formula to solve for "e":
Crumpp
-
Yes the function can be inverted to define the aspect ratio as a function of e, but that wouldn't mean that aspect ratio is a function of e alone.
Correct.
Nobody is claiming it is either.
Crumpp
-
Boy, I've missed the fun! :)
Let me give it a shot at answering the question.
The answer is yes with a caveat of depending on how accurate you want to estimate e factor.
The way I understand that it works the way it does is because Prandtl's lifting line theory makes certain assumptions that ignores the differences between different planform shapes.
"In particular, the lifting-line model ignores the effect of chordwise distribution of vorticity on the downwash distribution since all the vorticity generated at a given spanwise location has been collasped to a single point....
....While lifting line theory is useful for approximating the performance of unswept, high-aspect-ratio wings once the chord distribution is fixed, the method is unable to account for any aerodynamic difference between wings due to different planform shapes."
(NASA-CR-191274 Analysis and Design of Planar and Non-Planar Wings for Induced Drag Minimization, 1992)
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930004275_1993004275.pdf (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930004275_1993004275.pdf)
From what it appears to me this was understood but the impact of which was largely ignored because the results are good enough.
"For studies requiring a higher degree of accuracy, lifting-surface theories have been used [as opposed to lifting line theory], but generally it has been found that the additional complexity of these methods has not sufficiently improved the predictions to warrant common use." (NACA Report 921 - Theoretical Symmetric Span Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form).
http://naca.larc.nasa.gov/reports/1948/naca-report-921/naca-report-921.pdf (http://naca.larc.nasa.gov/reports/1948/naca-report-921/naca-report-921.pdf)
By the way, I've seen in from other sources but explicitly in my copy of Raymer (Aircraft Design: A Conceptual Approach) he gives as one of the methods of estimating e for straight wing aircraft as:
e = 1.78(1- .045*AR^.68) - .64
The basis for this equation comes from a report I don't have access to which appears to have derived the equation basis empirical results from actual aircraft. ("Subsonic Drag Estimation Methods" Cavallo, B., U.S. Naval Air Development Center Rept NADC-AW-6604, 1966.)
Anway, that's my attempt at an answer to the question!
Tango, XO
412th FS Braunco Mustangs
-
dtango,
Ah, now I see from where the formula in the Zigrat's spreadsheet comes from:
e=1*1,78*(1-0,0455*AR^0,68)-0,64
Give me your e-mail.
Regarding accuracy, these AR based formulas seem to give rather high e factor. When I have moretime I'll study couple more drag polars.
gripen
-
gripen:
dytong@nstci.com :)
Regarding the accuracy of e prediction using the above formula, I'm certainly no expert on the matter and the thought of using navier-stokes or panel methods for estimating this stuff makes my head swim.
Whatever the case, this is what Raymer states:
"Numerous estimation methods for e have been developed over the years, such as those by Glauert and Weissinger. These tend to produce results higher than the e values of real aircraft. More realistic estimation equations based upon actual aircraft (Ref 45.) are presented below."
Ref 45 is the NADC document and immediately following the above text are formulas - one for e of straight wing aircraft and one for e of swept wings.
I've seen the e formulas in question listed in several other aerodynamics text as well but never really understood how the equations were derived.
I'd be very curious to get a look at the NADC document to see how they boiled it down.
Tango, XO
412th FS Braunco Mustangs
-
Originally posted by gripen
Well, if we have no idea about taper ratio, there is no way we can estimate the e factor with the aspect ratio only. However Badboy's poin't is quite sensible; most rectangular winged WWII fighters had quite similar taper ratio so in many cases generalized formula do quite well.
I have calculated some e factor data from the drag polars I have. These are e factors for whole airplane as Oswald and Perkin&Hage handle it:
Bf 109G, this is from the VL (Finnish State aircraft factory) drag polar:
Cl 0,1 => 0,741658161
Cl 0,2 => 0,798708788
Cl 0,4 => 0,769126981
Cl 0,6 => 0,759747384
Cl 0,8 => 0,787352739
Cl 1,0 => 0,769126981
Cl 1,2 => 0,713350597
Cl 1,4 => 0,619327448
The average is about 0,75
I have also a polar from Messerschmitt AG (I had to approximate Cd0):
Cl 0,1 => 1,038321425 (not in average)
Cl 0,2 => 0,83065714
Cl 0,3 => 0,778741069
Cl 0,4 => 0,755142854
Cl 0,5 => 0,70156853
Cl 0,6 => 0,692214283
Cl 0,7 => 0,696955477
Cl 0,8 => 0,692214283
Cl 0,9 => 0,7067566
Cl 1,0 => 0,711179058
The average 0,73
Mustang, RAE wind tunnel data:
Cl 0,6 => 0,789174119
Cl 0,4 => 0,814631349
Cl 0,2 => 0,717431018
The average is about 0,77
I also tried to calculate from the NACA 916 but there is so much varition that sensible analysis seems to be quite impossible; the e factors varied from 0,5 to 1,3 or something.
Regarding prizes; well I think the answer is not possible but in certain conditions it might be possible as Badboy pointed out.
So I think I'll send the prize to Angus and Blogs and Schutt will receive the Bonus, so post your adress and I'll send the docs. If Badboy is listening I can send both documents to you too.
gripen
Am I the only one that has absolutley no idea what all that stuff is? :confused: Im lost LOL
-
madness: hehe :)
Those are values of Oswald efficiency relative to a given angle of attack (lift coefficient).
Gripen's point is that looking at the drag polars the values for oswald efficiency doesn't quite square with NACA report 916 nor with the method of estimating oswald efficiency factor using an equation based primarily on aircraft wing aspect ratio.
Tango, XO
412th FS Braunco Mustangs
-
The impression I get from the handful of sources I looked at is that the adjustment for departures from an elliptical lift distribution doesn't vary all that much among monoplanes of the 30s and 40s, but it does vary. Now I realize I could have that wrong.
So in terms of absolute efficiency of a wing, relying soley on aspect ratio to calculate induced drag may not be too bad. Problem is that when you compare two planes, it's the relative differences that matter. It is then very hard to say whether a difference in this parameter between two wings is going to decide whether one plane will do something better than another.
But this is largely a distraction from the main point I made at the start. The wing cross section, or profile, is the primary determinant of parasitic drag. In the total drag calculations we've been dragging around, we are treating this as a constant (and ignoring it). This parameter is going to vary across wings even with the same aspect ratio and it will affect the relationship between lift and total drag.
-Blogs
Originally posted by dtango
Boy, I've missed the fun! :)
Let me give it a shot at answering the question.
The answer is yes with a caveat of depending on how accurate you want to estimate e factor.
The way I understand that it works the way it does is because Prandtl's lifting line theory makes certain assumptions that ignores the differences between different planform shapes.
"In particular, the lifting-line model ignores the effect of chordwise distribution of vorticity on the downwash distribution since all the vorticity generated at a given spanwise location has been collasped to a single point....
....While lifting line theory is useful for approximating the performance of unswept, high-aspect-ratio wings once the chord distribution is fixed, the method is unable to account for any aerodynamic difference between wings due to different planform shapes."
(NASA-CR-191274 Analysis and Design of Planar and Non-Planar Wings for Induced Drag Minimization, 1992)
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930004275_1993004275.pdf (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930004275_1993004275.pdf)
From what it appears to me this was understood but the impact of which was largely ignored because the results are good enough.
"For studies requiring a higher degree of accuracy, lifting-surface theories have been used [as opposed to lifting line theory], but generally it has been found that the additional complexity of these methods has not sufficiently improved the predictions to warrant common use." (NACA Report 921 - Theoretical Symmetric Span Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form).
http://naca.larc.nasa.gov/reports/1948/naca-report-921/naca-report-921.pdf (http://naca.larc.nasa.gov/reports/1948/naca-report-921/naca-report-921.pdf)
By the way, I've seen in from other sources but explicitly in my copy of Raymer (Aircraft Design: A Conceptual Approach) he gives as one of the methods of estimating e for straight wing aircraft as:
e = 1.78(1- .045*AR^.68) - .64
The basis for this equation comes from a report I don't have access to which appears to have derived the equation basis empirical results from actual aircraft. ("Subsonic Drag Estimation Methods" Cavallo, B., U.S. Naval Air Development Center Rept NADC-AW-6604, 1966.)
Anway, that's my attempt at an answer to the question!
Tango, XO
412th FS Braunco Mustangs
-
The basis for this equation comes from a report I don't have access to which appears to have derived the equation basis empirical results from actual aircraft. ("Subsonic Drag Estimation Methods" Cavallo, B., U.S. Naval Air Development Center Rept NADC-AW-6604, 1966.)
A friend of mine is a Test pilot for the U.S. Naval Air Development Center. He is coming to visit soon. In fact he hooked me up with one his professor's in college who worked on the V2 program in WWII. I will be interviewing his professor and some other contacts who worked at Rechlin for my book.
I will see if he can't dig up a copy of that report and email it to me.
Crumpp
-
This parameter is going to vary across wings even with the same aspect ratio and it will affect the relationship between lift and total drag.
Exactly, if I understand it correctly it will vary with the twist of the wing, shape of the leading edge, and the taper ratio among other things.
The thought of calculating all that seems rather daunting!
Seems though that the more detailed the analysis, the more Davids Lednicer's conclusions seem correct.
Crumpp
-
Originally posted by Crumpp
Seems though that the more detailed the analysis, the more Davids Lednicer's conclusions seem correct.
Hm... I wonder what Lednicer's conclusions has to do with the subject of this thread? He calculated the span loading and concluded that the loading distribution of the Spitfire is "not elliptical, though it is probably the most optimum of the three aircraft from the induced drag standpoint". He did not determine Oswald's efficiency for the wing nor for the whole aircraft.
gripen
-
That's what's so fascinating with aerodynamics.
It's so "twisty".
Nothing is just perfect in all matters.
-
Hm... I wonder what Lednicer's conclusions has to do with the subject of this thread? He calculated the span loading and concluded that the loading distribution of the Spitfire is "not elliptical, though it is probably the most optimum of the three aircraft from the induced drag standpoint". He did not determine Oswald's efficiency for the wing nor for the whole aircraft.
His comparison of drag is totally different from the one you keep calculating. The more detailed you make the calculations the more his conclusions are correct.
In other words:
When you stop assuming values and actually calculate them.
It is not a ding on your intelligence Gripen. It's a matter of knowing the correct values and the formulas to use.
Obviously the formulas and data can be manipulated to produce wrong conclusions and still be correct.
If we factor in history it backs up Lednicers conclusions as well. If Merlin Power spits were capable of converting airspeed to altitude on the same efficiency level as the FW-190 then the RAF would have recommended that they do so. The R.A.E. had plenty of FW-190A's to test.
Historical facts are:
1. Merlin Spitfires never fought FW-190's in the vertical. They remained in the Horizontal where they had real advantages.
2. FW-190A's did not fight Spitfires by making clumsy attacks and extending beyond visual range. They got close and fought in the vertical where they had some real advantages.
FW-190's could only do this if they had some aerodynamically advantages in the vertical beyond simple mass. Look at the P47D11, a much heavier plane that the FW-190A easily out zoomed. In fact the only option a P47D11 had in fighting an FW-190A5 was to dive away as recommended by the USAAF. This was the same FW-190A5 used in the US Navy test's with the same problems. Only difference between a P47D4 and a P47D11 is the D11 is fitted with a bomb rack which further reduces it's performance. You can read the conclusions here:
http://prodocs.netfirms.com/
Crumpp
-
Crumpp,
This thread is about estmation of the Oswald's efficiency. Why are you bringing in issues which has nothing to do with it?
gripen
-
This thread is about estmation of the Oswald's efficiency. Why are you bringing in issues which has nothing to do with it?
To answer this:
Hm... I wonder what Lednicer's conclusions has to do with the subject of this thread? He calculated the span loading and concluded that the loading distribution of the Spitfire is "not elliptical, though it is probably the most optimum of the three aircraft from the induced drag standpoint". He did not determine Oswald's efficiency for the wing nor for the whole aircraft.
Because you keep skirting around the issue in both threads.
Crumpp
-
Crumpp,
Well, you are presenting quite a logic here; first you bring Lednicer in this thread and after I point out that Lednicer said quite little about the span loading and did not determine Oswald's efficiency at all, you blame me for skirting around the issue.
gripen
-
after I point out that Lednicer said quite little about the span loading and did not determine Oswald's efficiency at all
Actually he says quite a bit about span loading and how the wing twist of the aircraft effect the wet lifting surfaces. It is covered in the first part of the article and is quite informative. That is the whole reason the Spitfire does not develop a truly elliptical wing.
He mentions the "e" factor and states it is better than the other fighters but not by much. Not nearly as large a gap as the "guestimates".
e = 1.78(1- .045*AR^.68) - .64
Is the formula in Zigrat's spreadsheet. If you find another formula you think will be "better" let me know and I will recalculate the figures.
Crumpp
-
Crumpp,
I wonder what is your point now, basicly you are repeating what I said above and actually Tango allready claimed that formula above too (as well as me).
Have you tested that formula if it gives similar values as the drag polars?
gripen
-
No I haven't Gripen.
What is the data you are using for the Bf-109G and I will run it through.
Crumpp
-
Originally posted by Crumpp
Is the formula in Zigrat's spreadsheet. If you find another formula you think will be "better" let me know and I will recalculate the figures.
Crumpp
The empirical formulae used by zigrat appears to have been derived some 20 years after WWII ended so it is possible that the data used to produce it might not be as representative as this formulae
e = 1/(0.9676+0.033*AR)
Which was derived in 1935 using data current at that time, and was quite possibly used in the design of fighters that saw service during the war. Here is a comparison of the values from each formulae, notice particularly the values for AR of 1 and 2, because that's a clue as to which formulae might be more appropriate. However, I will admit that given the sacrifices already made in the assumptions used to derive these equations, this is a bit like splitting hairs :)
AR___NADC___Wood
1____1.06____1.00
2____1.01____0.97
3____0.97____0.94
4____0.93____0.91
5____0.90____0.88
6____0.87____0.86
7____0.84____0.83
8____0.81____0.81
9____0.78____0.79
10___0.75____0.77
11___0.73____0.75
12___0.70____0.73
Hope that helps...
Badboy
-
The data is VL and MTT data as claimed above. The first is from the VL archives and later is from the NASM.
gripen
-
e = 1/(0.9676+0.033*AR)
It changes both the Spitfires and the FW-190's "e" factor in the exact same proportion as:
e = 1.78(1- .045*AR^.68) - .64
Different values with the same spread.
I will run some test calculations but I think the conclusions will be exactly the same suing either formula.
Crumpp
-
Tango!? Is that you bud? Are ya the same tango from ww2ol that flew with 8th af?
-
madness:
Sorry not the same tango. I've never played WW2OL.
Tango, XO
412th FS Braunco Mustangs
-
Crumpp,
You don't seem to understand how to tests these formulas; you need a drag polar of a certain plane to make comparison, see Perkins&Hage p. 90-94.
gripen
-
Ahh ok, thought you were somebody else :lol oh well. Thanks for explaining that jibberish :D .
S!
-
You mean this one?
http://www.wiley.com/WileyCDA/WileyTitle/productCd-047168046X.html
Crumpp
-
Yep, that's the reference regarding WWII fighter performance estimation.
gripen
-
Got it on order. Should be here in a couple of days.
Crumpp
-
I can't quite follow all of this, but do you guys have any suggestions on books or papers where these formula are explained?
[edit]Other than what has already been linked to. :)
-
Starbird:
Here are three of my favorite sites that explain a lot of the fundamentals:
Aerodynamics Index (http://142.26.194.131/aerodynamics1/)
US Centennial of Flight: Theory of Flight (http://www.centennialofflight.gov/essay_cat/9.htm)
NASA Beginner's Guide to Aerodynamics (http://www.grc.nasa.gov/WWW/K-12/airplane/short.html)
There are a lot of books out there. I did come across one that was really helpful in terms of explaining the concepts at fundamental and intermediate levels without getting heavy into the math but I can't remember the name of it. I borrowed it from my uncle awhile back and it was a short digestable book geared for pilots. I'll see if I can dig it up.
Hope that is helpful!
Tango, XO
412th FS Braunco Mustangs
-
From the first reference we have:
"From our earlier discussions, we know that up wash increases and therefore induced drag increases when CL2 increases. We also predicted that induced drag would be inversely proportional to the aspect ratio (AR).
A full development of the equation for Cdi (coefficient of induced drag) is beyond the scope of this course, however the equation is:
Cdi = CL²/ (pi*e*AR)
In the equation e is a number known as the Oswald efficiency factor. It is a number less than or equal to one. The Oswald efficiency number would equal 1 only for an elliptical wing with no fuselage or other appendages to affect the up wash. In the real world of course, most aircraft do not have elliptical wings and other factors such as the fuselage may also contribute to up wash. Therefore, e is always less than 1, even on a Spitfire. Usually e’s value is between .7 and .9. Low values will be for aircraft with boxy square wings and large fuselages. Aircraft with long tapered wings (like gliders and most airliners) will have high values for e."
Nuff said
-blogs
Originally posted by dtango
Starbird:
Here are three of my favorite sites that explain a lot of the fundamentals:
Aerodynamics Index (http://142.26.194.131/aerodynamics1/)
US Centennial of Flight: Theory of Flight (http://www.centennialofflight.gov/essay_cat/9.htm)
NASA Beginner's Guide to Aerodynamics (http://www.grc.nasa.gov/WWW/K-12/airplane/short.html)
...
Tango, XO
412th FS Braunco Mustangs
-
Here is the e factor estimate for the Spitfire, this is based on the RAE wind tunnel data:
Cl 0,1 0,565121862
Cl 0,2 0,869418249
Cl 0,3 0,847682793
Cl 0,4 0,904194979
Cl 0,5 0,94186977
Cl 0,6 0,847682793
The average is 0,88
I had no wind tunnel data to calculate e factor of the Fw 190. But Lednicer calculated span loadings and here is estimated e factors calculated by taking 8 samples (semispan fractions 0,2-0,9) and measuring distance to the elliptical span loading:
Spitfire: 0,8684
Fw 190: 0,7857
P-51: 0,7965
The values of the Spitfire and P-51 are suprisingly close to the values which are calculated from the drag data above. So it is quite safe to say that the value for the Fw 190 is in the right ballpark.
Below is an example of the drag polar. It is the drag polar of the Bf 109G from the Messerschmitt AG.
Blogs, could you give your e-mail address.
gripen
(http://personal.inet.fi/koti/soon.moro/bf109g_Cd.jpg)
-
Here is the e factor estimate for the Spitfire, this is based on the RAE wind tunnel data:
Interesting, because Zigrats spread sheet calculates the "e" factor for the Spitfire to be:
Spitfire = 0.88552405
Which is EXACTLY what you say the average should be!
The average is 0,88
It calculates the FW-190 "e" factor to be:
FW-190 = 0.865317063
Which is nothing like:
Fw 190: 0,7857
And therefore probably nothing like:
So it is quite safe to say that the value for the Fw 190 is in the right ballpark.
What is going on?
Either the math only works for the Spitfire by some fluke of GOD.
or
Something is not right with your calculations.
Crumpp
-
Woods formula for the Spitfire lends:
Spitfire = 0.873125707
Again within your ballpark.
FW-190A8 = 0.857305852
Again NOTHING close to your calculations.
It backs up David Lednicers analysis though.
Crumpp
-
Woods formula for the Spitfire lends:
Spitfire = 0.873125707
Again within your ballpark.
FW-190A8 = 0.857305852
Again NOTHING close to your calculations.
It backs up David Lednicers analysis though.
Firstly Wood's and the NADC formula's for e are all approximations based mainly on AR. If you're expecting values to match up perfectly with everything you're expecting a bit much.
2ndly I'm not sure how any of this has anything to do with Dave Lednicer's analysis since (1) the flat plat drag values he quotes aren't things he calculated using VSAERO, the CFD s/w he used that I'm aware of and (2) he doesn't mention any specific analysis of oswald efficiency factor which is a specific component of lift dependent drag breakdown.
Seriously crumpp, where are you going with all this? I had enough of the last thread because of the "I'm right no matter what" routine. I hope we're not going down that path.
Tango, XO
412th FS Braunco Mustangs
-
Crumpp,
Seems that you don't fully understand the issue here. All my values are based on the documentation on the discused planes and the calculated e factors are following:
Bf 109G: 0,73-0,75
P-51: 0,77-0,8
Spitfire: 0,87-0,88
Fw 190: 0,79
The aspect ratios are following:
Bf 109G: 6,13
P-51: 5,81
Spitfire: 5,61
Fw 190: 6,02
The taper ratios:
Bf 109G: 0,44 (0,5 without wing tips)
P-51: 0,46
Spitfire: elliptical
Fw 190: 0,44
The washout:
Bf 109G: 0
P-51: 2 deg
Spitfire: 2,25 deg
Fw 190: 2 deg
If we look these numbers and compare them to generalized formulas (like Wood's or the one in the Zigrat's sheet), we can see that except the case of the Spitfire, generalized formulas seem to give somewhat higher values (around 10%, and in the case of the Spitfire the accuracy seem to be more or less accidental). We can also see that this very limited data set supports assumption that the e factor decreases when the aspect ratio increases. In addition we can also see that taper ratios are somewhat lower than assumed in the Wood's formula and except the case of the Bf 109, the planes have some washout (I don't know if the generalized formulas assumed washout).
Anyway, this is by far too limited data set; it would most interesting to have drag polars of the planes with higher aspect ratio (like the Ta 152H and the P-38) and the P-47 would be also interesting due to elliptical wing shape.
gripen
-
Firstly Wood's and the NADC formula's for e are all approximations based mainly on AR. If you're expecting values to match up perfectly with everything you're expecting a bit much.
If we look these numbers and compare them to generalized formulas (like Wood's or the one in the Zigrat's sheet), we can see that except the case of the Spitfire, generalized formulas seem to give somewhat higher values (around 10%, and in the case of the Spitfire the accuracy seem to be more or less accidental).
I was not referring to Lednicers flat plate comparisons. I was referring to his discussion on the elliptical shape of the lifting wet area.
The Spitfires physical elliptical shape was an attempt to produce an elliptical wet lifting area, correct. The problem he states with that is the stall characteristics. To eliminate the problematic stall of a truly elliptical wing, the Spitfires wings are twisted just like most fighters. That twist according to him, removes the major advantages and eliminates the huge advantage everyone thinks the spitfire had in "e" factor.
He states that of the loading distribution of the spitfire is NOT elliptical. Of the three it probably has the best "e" factor but it does not have a huge advantage.
case of the Spitfire the accuracy seem to be more or less accidental
That seems highly suspicious. Especially since all the other data lines up perfectly with Lednicer's conclusions. He used the actual German drag plots as noted in his references.
Crumpp
-
...according to him, removes the major advantages and eliminates the huge advantage everyone thinks the spitfire had in "e" factor. ....He states that....Of the three it probably has the best "e" factor but it does not have a huge advantage.
Firstly, Dave Lednicier doesn't say anything about the quantitative advantage of one aircraft over another based on the lift distribution so it's unfair to claim that he says anything about the magnitude of the advantage.
2ndly most students of aerodynamics are aware of the famous elliptical wings of the Spitfire, what they were trying to achieve with it and HOW it didn't meet the ideal. This is nothing new. Dave Lednicer points out specifics of WHY it didn't. I'm not sure where you get the notion about what "everyone" believes about the Spitfire.
3rdly I'm not sure why we're even talking about this since this thread was a discussion regarding estimating span efficiency factor using AR as the prime variable.
Tango, XO
412th FS Braunco Mustangs
-
Originally posted by Crumpp
He states that of the loading distribution of the spitfire is NOT elliptical. Of the three it probably has the best "e" factor but it does not have a huge advantage.
I wonder what you are trying to prove? If you look my numbers, you will note that the e factor of the Spitfire is not 1. My numbers also show that the Spitfire has best e factor of the three. Those Generalized models give the best e factor for the Spitfire just because it has the lowest aspect ratio of the three and therefore relative difference between planes is wrong.
gripen
-
I wonder what you are trying to prove? If you look my numbers, you will note that the e factor of the Spitfire is not 1. My numbers also show that the Spitfire has best e factor of the three. Those Generalized models give the best e factor for the Spitfire just because it has the lowest aspect ratio of the three and therefore relative difference between planes is wrong.
Knowing your prejudice for the Spitfire, Gripen. Just look at the merry -go-round we just went through in the last thread.
http://www.hitechcreations.com/forums/showthread.php?s=&postid=1391720#post1391720
First you make a bunch of assumptions that show the FW-190 way behind the spitfire in flat plate areas that are proved wrong in the math. Even using ALL the values you claimed!
Then you claim it is only JL 165 that has a drag disadvantage.
When that is proved wrong your story changes again.
Then it changes and is the Spitfire Mk VIII which has the drag advantage.
In fact, when the numbers are run ONLY ONE Spitfire Mk IX had the advantage. The one with an new style of air intake. And of course the Spitfire Mk VIII which is a at totally different plane.
None of the Spitfire's had a parasitic drag advantage.
Now your making calculations off of assumptions AGAIN.
Anyway, this is by far too limited data set; it would most interesting to have drag polars of the planes with higher aspect ratio (like the Ta 152H and the P-38) and the P-47 would be also interesting due to elliptical wing shape.
In fact lets review what the Aeronautical engineering instructor says about taper ratio:
However, the formulae presented by Professor Wood is for a taper ratio of 0.57 which yields an almost elliptical lift distribution, which is why it fits the experimental data so closely over a wide range of examples. Most WWII aircraft had a taper ratio close to that value (Me109 was approx 0.52) because the designers knew about the benefits of elliptical lift distribution and that it could be achieved quite closely with a wing of that taper, that’s why a formulae that only includes aspect ratio could still be of such good practical use.
Spitfire: elliptical
Your making the assumption in your calculations that the Spitfire gains benefits for it's elliptical shaped wing. The entire point is that when tested, the wing was not elliptical due to the twist in the outboard portion.
Your simply changing the formula to one that better suits your plane.
You explain away the "fluke" of the NACD and woods formulas for Oswald's efficiency coming so close the your favorite plane, the Spitfire, by declaring:
If we look these numbers and compare them to generalized formulas (like Wood's or the one in the Zigrat's sheet), we can see that except the case of the Spitfire,
I say baloney. You're playing the shell game again.
The values of the Spitfire and P-51 are suprisingly close to the values which are calculated from the drag data above.
There is no drag data on the P51 accept a number you calculated. Only actual data you present is the Bf-109G, which is not the A/C in question.
Crumpp
-
Originally posted by Crumpp
First you make a bunch of assumptions that show the FW-190 way behind the spitfire in flat plate areas that are proved wrong in the math.
Hm... My assumption regarding the e factor was 0,9 for the Spitfire (calculated 0,87-0,88) and 0,8 for the other two (calculated 0,77-0,80) so with my assumptions the errors were small and overall smaller than with generalized model.
Regarding data, the Mustang data is available from the PRO as DSIR 23/14544 and the Spitfire data as DSIR 23/12576.
gripen
-
Hm... My assumption regarding the e factor was 0,9 for the Spitfire (calculated 0,87-0,88) and 0,8 for the other two (calculated 0,77-0,80) so with my assumptions the errors were small and overall smaller than with generalized model.
No, not really. Your third or so assumptions might have been in that ballpark but not your original ones. After you were confronted with calculations. Pretty sure that is not the right ball park either.
overall smaller than with generalized model
Which is much more accurate than your assumptions.
Right now your trying to argue that the 190's design team, using a wing designed after the Spitfire, in a time when designers knew about the benefits of elliptical distribution, did not plan for it. Seems kinda silly to me.
Crumpp
-
I was under the impression that Heinkel designed their pre-war commercial plane (can't remember the model number) with an elliptical planform in mind because of prandtls work. And I thought this had influenced the designers at supermarine too.
So wouldn't this be old news by 1940?
-Blogs
Originally posted by Crumpp
...
Right now your trying to argue that the 190's design team, using a wing designed after the Spitfire, in a time when designers knew about the benefits of elliptical distribution, did not plan for it. Seems kinda silly to me.
Crumpp [/B]
-
I was under the impression that Heinkel designed their pre-war commercial plane (can't remember the model number) with an elliptical planform in mind because of prandtls work. And I thought this had influenced the designers at supermarine too.
It was old news by then. That is the point. Once you understand the physical shape has little to do with the wet area shape you can see planes like this one:
http://www.anycities.com/user/j22/j22/index.htm
Have an "e" factor of .95
http://www.anycities.com/user/j22/j22/aero.htm
And that just because the spitfire had a physically elliptical shape has nothing to do with it's "e" factor given the wingtip twist.
Lednicer explains the elliptical shape was more an attempt to better deal with the problem's of wing mounted landing gear than induced drag. That allowed the spitfire wing to be relatively thin. We all know Willey went around it by putting the landing gear in the Fuselage. That has it's own set of unique problems though.
Crumpp
-
Originally posted by Crumpp
Which is much more accurate than your assumptions.
Let's see:
NADC formula (As Zigrat's sheet) in the best case:
Spitfire: about right
P-51: about 7 units of
Fw 190: about 8 units of
=> average error is about 5 units
My assumption in the worst case:
Spitfire: 3 units of
P-51: 5 units of (0,75)
Fw 190: 1 unit of
=> average error is 3 units
Even when I used 0,8 for all three, the average error was 4 units ie less than the generalized formula in the best case.
gripen
-
Even when I used 0,8 for all three, the average error was 4 units ie less than the generalized formula in the best case.
The Generalized formulas put the "e" factors at much closer than assuming the same value for all.
They also put the Spitfire in the lead for "e" factor.
The Mustand and the FW-190 are very close with the Mustang having a slight lead.
Spitfire = .871
FW-190 = .857
P51 = .861
Your calculations do not account for nearly the number of factors the theory behind the "general" formulas do, especially the more modern NACD formula.
The designers of these wings were well aware of the benefits of elliptical distrubtion and designed their wings accordingly.
Low aspect ratio wings have nearly elliptic distributions of lift for a wide range of taper ratios and sweep angles. It takes a great deal of twist to change the distribution. Very high aspect ratio wings are quite sensitive, however and it is quite easy to depart from elliptic loading by picking a twist or taper ratio that is not quite right.
Note that many of these effects are similar and by combining the right twist and taper and sweep, we can achieve desirable distributions of lift and lift coefficient.
http://www.desktopaero.com/appliedaero/wingdesign/geomnldistn.html
Ideally, this drag contribution should be estimated for the individual airplane components, with factors such as the influence of wing leading edge geometry, wing camber, wing thickness ratio, wing sweep, pylon interference, fuselage upsweep, tail induced drag, power effects, etc. taken into account.
http://adg.stanford.edu/aa241/drag/induceddrag.html
All of these things are considered in the theory behind the "general" formulas.
You assume the Spitfire has perfect theoritical elliptical form. In fact it does not.
http://www.onemetre.net/Download/Downwash/Downwash.htm
Wing sweep is chosen almost exclusively for its desirable effect on transonic wave drag. (Sometimes for other reasons such as a c.g. problem or to move winglets back for greater directional stability.)
http://adg.stanford.edu/aa241/wingdesign/wingparams.html
None of your calculations account for the sweep of both the 190's wing and the P51's.
Crumpp
-
Originally posted by Crumpp
The Generalized formulas put the "e" factors at much closer than assuming the same value for all.
As noted couple times above, these generalized formulas seem to give rather high e factors. Therefore even a single value in the right ballpark is more accurate as seen above.
Originally posted by Crumpp
You assume the Spitfire has perfect theoritical elliptical form.
Only you have made such statement in these discussions.
Originally posted by Crumpp
None of your calculations account for the sweep of both the 190's wing and the P51's.
This argument is laughable and speaks a lot about your knowledge.
gripen
-
As noted couple times above, these generalized formulas seem to give rather high e factors. Therefore even a single value in the right ballpark is more accurate as seen above.
No, your calculations don't have all the data.
At least according to you:
We can also see that this very limited data set supports assumption
Means - We can just guess because we don't have the facts.
This argument is laughable and speaks a lot about your knowledge.
Not really, I was refering to the fact that wingsweep:
We can use this fact to design a wing which can fly at a high speed with a pressure distribution associated with a lower speed.
http://www.desktopaero.com/appliedaero/potential3d/sweeptheory.html
Adds benefits by reducing the CL and thereby reducing the CDi while going the same speed as an unswept wing like the Spitfire.
Crumpp
-
For Crumpp:
I've not been reading the whole thread, but glancing through it I came across some stuff about the Spitfire's elliptical form.
I have an article about this somewhere, - I'll gladly scan it and send it to you. But shooting from memory:
1:The Elliptical form has about 10% less induced drag than a square wing with the same aspect ratio, or at least 5% less than any other wingform
2:A completely elliptical wing stallls entirely, since the lift distribution is equal. (It stalls LATER than any other wing of the same size, but when it stalls, it stalls!)
3: The Spitfire Wing was not completely elliptical, - but however very very close. I am not entirely sure about this, but I belive that the ellipse got "watered out" at the wingtips. Longer span then and at stall, the aircraft hanging on the tips, however a worse rollrate as a payback. I would not risk my neck on this particular item, but it's worth having a look into and I am sure that you guys will
:)
-
Oh, BTW, there is actually one type of an aerobatic plane about which has an elliptical wing. It is favoured for it's ability to go through an aerobatic sequence with the minimal loss of energy, - being somewhat underpowered. Just can't find the book about it right now.
Anyway, Gripen and Crumpp:
As much as I like you both, how about a slightly more friendly tone in your salvoes?
Gripen?
-
Originally posted by Crumpp
No, your calculations don't have all the data.
At least according to you:
That means that e factor data for 4 planes is a limited set, large set would be say 40 planes. The calculations are based on the documented Cl /Cd data.
Originally posted by Crumpp
Adds benefits by reducing the CL and thereby reducing the CDi while going the same speed as an unswept wing like the Spitfire.
Why don't you just read your source:
"Furthermore, at a given angle of attack, the lift is reduced."
Basicly the drag for given lift increases with the sweep. That should be easy to understand.
gripen
-
but I belive that the ellipse got "watered out" at the wingtips.
Thanks Angus. Appreciate your input bro. The wingtips are most important part of the wing we are discussing. Your absolutely right and that is the whole point. The Spitfire had about the best efficiency factor of the 3 planes we are discussing. It's lead is not that large though because of it's "watered out" ellipse.
It's Aspect Ratio and Taper Ratio are more responsible for its lead in "e" factor than it's "elliptical" wings.
"Furthermore, at a given angle of attack, the lift is reduced."
No I understand it Gripen. What you are leaving out is the fact that SPEED is increased. A straight wing fighter going the same speed will work harder to achieve that speed.
At the same speed the swept wings CL will be lowered proportional to the Cos of the angle of sweep thereby lowering it's CDi accordingly.
That is why swept wing fighters have such high approach speeds and such sharp AoA on approach.
Furthermore Oswald's efficiency factor is influenced by sweep, parasitic drag, and AR.
The figure below shows a typical variation in e with aspect ratio, sweep, and CDp
Again this whole thing comes down too:
Right now your trying to argue that the 190's design team, using a wing designed after the Spitfire, in a time when designers knew about the benefits of elliptical distribution, did not plan for it. Seems kinda silly to me.
Crumpp
-
That means that e factor data for 4 planes is a limited set, large set would be say 40 planes. The calculations are based on the documented Cl /Cd data.
Using what calculations? We have already seen the varying results of different formulas.
Crumpp
-
Oh, BTW, there is actually one type of an aerobatic plane about which has an elliptical wing.
The only well known modern aerobatic aircraft I can think of is the CAP 10B (http://www.airliners.net/open.file?id=621258&WxsIERv=TXVkcnkgQ0FQLTEwQg%3D%3D&WdsYXMg=VW50aXRsZWQ%3D&QtODMg=QmVybiAtIEJlbHAgKEJlbHBtb29zKSAoQlJOIC8gTFNaQik%3D&ERDLTkt=U3dpdHplcmxhbmQ%3D&ktODMp=SnVuZSAyNywgMjAwNA%3D%3D&BP=1&WNEb25u=QnJ1bm8gQWx0aGF1cw%3D%3D&xsIERvdWdsY=SEItU0JE&MgTUQtODMgKE=Q2Fub24gMTBEIC8gSVNPMTAwIC8gMTAwLTQwMG1tIC8gRjYuNyAvIDEvMzUwcw%3D%3D&YXMgTUQtODMgKERD=MjM5&NEb25uZWxs=MjAwNC0wNy0xNQ%3D%3D&ODJ9dvCE=&O89Dcjdg=MjY4&static=yes).
-
Originally posted by Crumpp
At the same speed the swept wings CL will be lowered proportional to the Cos of the angle of sweep thereby lowering it's CDi accordingly.
So above you actually say that the lift coefficient of the swept wing will be lowered at given speed, that means that the angle of attack must be increased to reach given lift at this speed if compared to the straight wing and that means higher induced drag at given speed. Great logic.
Originally posted by Crumpp
Using what calculations? We have already seen the varying results of different formulas.
If you know the Cd at the given Cl and AR, the e factor is easy to calculate. All what is needed is the drag polar of a given plane. If you can't figure out how, check Perkins&Hage or use solver.
Angus,
Please look other side's arguments a bit more closely.
gripen
-
Maybe slightly off-topic, but here's Mark Drelas explanation of wing sweep:
(http://yarchive.net/air/sweep.html )
Sweeping a wing makes sense only if you are up against the Mach number limit, and want to fly faster, as with a jetliner. It doesn't make sense if you want to fly higher, as with the U-2, or if Mach is of no concern, such as with a General Aviation aircraft.
The airfoils on a swept wing behave as though they were flying at a reduced speed, reduced Mach number, and reduced dynamic pressure.
effective speed = V cos(L)
effective Mach = M cos(L)
effective q = 0.5 rho V^2 [cos(L)]^2
where L is the sweep angle, and V and M are the airplane's speed and Mach.
Imagine a straight-wing airplane flying at its maximum Mach number. As you sweep the wing in flight from 0 to L degrees, the available lift drops by a factor of [cos(L)]^2, and the Mach compressibility effects on the wing's airfoils decrease (weaker shocks, etc.). You then increase the speed by a factor of 1/cos(L), so that the effective dynamic pressure and lift are increased back to their original levels. The effective Mach is also increased back to its original level. In effect, you haven't done anything to the wing's lift or compressibility effects, but the airplane is now flying faster!
In reality, this isn't a complete freebie, since the skin friction drag has increased by a factor of [1/cos(L)]^2 -- the wing skin friction isn't affected by sweep very much, and feels the full brunt of the real dynamic pressure increase, just like the rest of the airplane. So the overall L/D will typically decrease from the sweep. An airliner depends on the higher speed to more than compensate for the lower L/D and give better overall range (the product V x L/D is what appears in the range equation). And of course flying faster gives faster revenue stream for the airlines.
Sweep doesn't make sense on slower piston and turboprop airplanes. In general, if Mach number is not a speed-limiting factor, it makes more sense to get more speed by reducing the wing area.
-
If you know the Cd at the given Cl and AR, the e factor is easy to calculate. All what is needed is the drag polar of a given plane. If you can't figure out how, check Perkins&Hage or use solver.
I know that Gripen.
We have already seen that their is more than one formula for every parameter in aeronautics.
This is exactly what the NASA aeronautical engineer cautioned about when using data from WWII documents. You have to know the formula used and the exact conditions of flight it was calculated for.
I have Perkins and Hage coming. Additionally I searching for the "Drag data for Aircraft" Focke-Wulf factory report and will get a copy.
So how did you get the CL and the drag for the FW-190?
Crumpp
-
Originally posted by Crumpp
We have already seen that their is more than one formula for every parameter in aeronautics.
Yes that's the reason for this thread. Estimating Oswalds's efficiency from only one parameter (aspect ratio) seems to be quite unaccurate.
The Cl/Cd data like drag polar actually contains all the needed variables given the testing is properly done.
Originally posted by Crumpp
So how did you get the CL and the drag for the FW-190?
It's desbribed above; not with drag polar but taking samples from the span loading data set and comparing results to the drag polar data of the other planes. And the results supported each other well.
gripen
-
Yes that's the reason for this thread. Estimating Oswalds's efficiency from only one parameter (aspect ratio) seems to be quite unaccurate.
No actually it is quite accurate and has the Spitfires "e" down exactly. I am sure if the FW-190 or P51 was your favourite plane they would be much closer as well.
It's desbribed above; not with drag polar but taking samples from the span loading data set and comparing results to the drag polar data of the other planes. And the results supported each other well.
Please point out the chart you are using to "guestimate" the FW-190's CL.
http://www.thetongsweb.net/AH/EAAjanuary1999.pdf
Crumpp
-
Gripen,
You list the Spitfires Taper Ratio as elliptical. It is not. The Spitfires TR should be around .67.
(http://www.onpoi.net/ah/pics/users/503_1095476250_downgraf.gif)
http://www.onemetre.net/Download/Downwash/Downwash.htm
According to this site the TR should be around .4 for elliptical performance:
A tapered wing concentrates the strength of the wing near to the center. Tapered wings approximate elliptical wings and are thus slightly more efficient than are straight wings. A taper ratio of 0.4 most closely approximates the elliptical. Highly tapered wings can suffer undesirable reynolds numbers causing the tip of the wing to stall earlier than the root at low speeds.
http://ourworld.compuserve.com/homepages/JHopkinson/wing2.html#WTR
Crumpp
-
Just heard back from Flugwerk.
They are building FW-190A8/N's and have restored several. According to them the "e" factor from FW documentation is:
I was checking my documents for the FW190 and the tip efficiency factor, or rather the Oswald Ellipse factor is about 1.15 or 0.87 depending on how you look at it. Basically it has about 15% more induced drag than a theoretical elliptical lift distribution.
Crumpp
-
Originally posted by Crumpp
You list the Spitfires Taper Ratio as elliptical. It is not. The Spitfires TR should be around .67.
I don't know if you understand at all what your source says. Basicly there is no exact taper ratio for the elliptical wing, one can calculate theoretical value but it has very little to do with the lift distribution and Oswald's efficiency.
Originally posted by Crumpp
According to this site the TR should be around .4 for elliptical performance:
Actually your source says that in the case of the tapered wing "A taper ratio 0,4 most closely approximates elliptical ".
That also rises an interesting question, who is right in this case. Your source or Badboy and professor Wood who claim that "taper ratio of 0.57 which yields an almost elliptical lift distribution"?
Who is right; Messerschmitt documentation which claims taper ratio 0,44 for the Bf 109G or the Aeronautical engineering instructor who claims taper ratio 0,52 for the Bf 109?
I was checking my documents for the FW190 and the tip efficiency factor, or rather the Oswald Ellipse factor is about 1.15 or 0.87 depending on how you look at it. Basically it has about 15% more induced drag than a theoretical elliptical lift distribution.
Who has actually written this?Are these values for the entire airframe (as I have calculated) or for the wing only? What is the Cl range of the measurements?
gripen
-
Who is right; Messerschmitt documentation
Who has actually written this?Are these values for the entire airframe (as I have calculated) or for the wing only? What is the Cl range of the measurements?
Focke-Wulf Gmbh., Bremen wrote it. It's the drag polars and it comes from this report.
(http://www.onpoi.net/ah/pics/users/503_1096577911_dragdatatitle.jpg)
Sent Pyro a copy.
Crumpp
-
Ah, so apparently you have the drag polars. Now just calculate e factor as described in the Perkins&Hage or the documentation might contain value of the K.
But actually I was asking which of your sources is right on taper ratio because you seem to use quite contradictory sources?
gripen
-
Now just calculate e factor as described in the Perkins&Hage or the documentation might contain value of the K.
It does.
The e factor averages out too:
0.87
You answer this:
But actually I was asking which of your sources is right on taper ratio because you seem to use quite contradictory sources?
here:
Basicly there is no exact taper ratio for the elliptical wing, one can calculate theoretical value
One thing is certain. The Spitfire did not have an truely elliptical wet lifting shape on it's wingtips as David Lednicers report concludes. Based on what TR you use will vary your calculations.
What the drag polars show us is that Zigrat does good work. His spreadsheet draws the correct conclusions in this case.
The Spitfire averages .88 and the FW-190 averages .87.
The FW-190 also always has less parasitic drag than the Spitfire.
Pyro and HTC have a copy of the report. I am sure their calculations will confirm this too.
Crumpp
-
Hey,
Do I get the prize?
Crumpp
-
My copy came today. Will take some time to digest.
One concern I have about all this debate is the acceptable margin of error. A lot of the engineering calculations substitute constants for other (sometimes unknown) equations. This was apparently good enough for the engineers at the time, but I worry the debate is running on such fine lines that it might affect the conclusions we reach.
-Blogs
Originally posted by gripen
So above you actually say that the lift coefficient of the swept wing will be lowered at given speed, that means that the angle of attack must be increased to reach given lift at this speed if compared to the straight wing and that means higher induced drag at given speed. Great logic.
If you know the Cd at the given Cl and AR, the e factor is easy to calculate. All what is needed is the drag polar of a given plane. If you can't figure out how, check Perkins&Hage or use solver.
Angus,
Please look other side's arguments a bit more closely.
gripen
-
One concern I have about all this debate is the acceptable margin of error. A lot of the engineering calculations substitute constants for other (sometimes unknown) equations. This was apparently good enough for the engineers at the time, but I worry the debate is running on such fine lines that it might affect the conclusions we reach.
I agree with you on this Joe. This debate did run on some fine lines.
Anyway, as far as I am concerned it is over. The question is answered.
Been pooring over my copy of Perkins&Hage, great info.
Crumpp
-
Crumpp,
I don't see the documentation here on how did you reach the e factor value 0,87 (or 1,15). According to Lednicer's span loading chart, the lift distribution of the Fw 190 was a bit worse than the lift distribution of the P-51. And the e factor of the P-51 seems to be around 0,75-0,8 so who is right, Lednicer or you?
What is the value of the K in the Fw drag polar?
And still no answer for contradictory taper ratio claims?
gripen
-
I don't see the documentation here on how did you reach the e factor value 0,87 (or 1,15).
No you don't. I will share it in the same manner you shared the polar plots you have in your possession.
And the e factor of the P-51 seems to be around 0,75-0,8 so who is right, Lednicer or you?
Ever consider a third possibility?
And still no answer for contradictory taper ratio claims?
It is answered, see above.
Crumpp
-
Here is the data Crumpp apparently don't want us to see:
(http://personal.inet.fi/koti/soon.moro/FwDrag.jpg)
It gives directly the value of the K at high speed and in the climb so the value of the e is very easy to calculate. At high speed the Cl is very low and probably not at linear stage of the Cd/Cl^2 curve so the climb speed value (1,24) is the one we are interested:
CwF = Cw0 + 1,24 * Ca^2
=>
Cd = Cd0(f) + 0,0678 * Cl^2
=>
0,0678 = 1 / (pi * AR * e)
=>
e = 0,78
At high speed the e value is 0,83.
Originally posted by Crumpp
No you don't. I will share it in the same manner you shared the polar plots you have in your possession.
Now you are creating double standards; I have given the source as well as place to get for every dataset I have used in this thread. In addition I have given directly the shape of the curves and the calculations. You have not given your data nor how did you reach your numbers.
Originally posted by Crumpp
Ever consider a third possibility?
All I see here is that even your own sources don't support your own arguments.
Originally posted by Crumpp
It is answered, see above.
I don't see your answer above. You have quoted two values, 0,4 and 0,52, which is right?
gripen
-
Your calculation for the high speed "e" factor of the Spitfire:
Cl 0,6 0,847682793
Your calculation for the high speed "e" factor of the FW-190
At high speed the e value is 0,83.
I think you don't have the whole report. If you compared the Spitfires climb "e" factor, I am pretty sure they would exhibit a similar spread.
Crumpp
-
Always bear in mind, that if calculations show that an aircraft should do something it doesn't, something could be wrong with the formula.
Edited due to rude wording originally.;)
-
Originally posted by Crumpp
Your calculation for the high speed "e" factor of the Spitfire:
...
Your calculation for the high speed "e" factor of the FW-190
It would help a lot if you could understand what we are talking about; the K value of the FW 190 at high speed is for the Cl value about 0,17 ie not for linear stage of the Cd/Cl^2 curve while the e value for the Spitfire you quoted is for Cl 0,6.
Basicly you continously quote numbers and link pages without understanding at all what do these mean.
Originally posted by Crumpp
I think you don't have the whole report.
Here is another example of your logic; using same logic I could say that you have only the headlines of the sheet because that is the part you linked above.
We need just values of the K for this thread and I linked them above, you did not.
gripen
-
It would help a lot if you could understand what we are talking about; the K value of the FW 190 at high speed is for the Cl value about 0,17 ie not for linear stage of the Cd/Cl^2 curve while the e value for the Spitfire you quoted is for Cl 0,6.
Gripen,
What you fail to understand is that I did not do the "e" factor calculations.
Some professional aeronautical engineering firms did. They did them for a FW-190A8 that is being restored to flight and will be flown exactly as it was in 1945. Minus Ammunition of course, that will have to be ballasted.
Those calculations are confirmed by another professional aeronautical engineering firm that is producing FW-190's. When calculated as per a fully loaded WWII Air Superiority fighter version.
In the other thread your willingness to manipulate data on the Spitfire was very evident. That puts any calculations you do under grave suspicion.
We need just values of the K for this thread and I linked them above, you did not.
So tell me what speed is that value for "K" calculated under? What speed is the spitfires? In order to compare both A/C must be under the SAME conditions of flight (altitude, speed, attitude, etc). Your putting up some nice calculations that are meaningless for comparision purposes and trying to compare. As it seen from the math, the efficiency factor does change under different flight conditions.
When I used the same logic and applied the "high speed" efficiency factor, using your "high speed" calculations, you call it into question. It is the SAME logic you are using in your comparisions.
Crumpp
-
Originally posted by Crumpp
What you fail to understand is that I did not do the "e" factor calculations.
So now you change your story again. Above you claim that e factor averages 0,87 which is actually nonsense because value of the K gives directly exact value of the e, there is no need to calculate average or fit the line as alternative method.
Originally posted by Crumpp
Some professional aeronautical engineering firms did. They did them for a FW-190A8 that is being restored to flight and will be flown exactly as it was in 1945.
Who actually did the calculation and how? Besides anyone can calculate e factor from the value of the K in less than five minutes as pointed out above. Apparently the numbers you got are wrong or the definition of the e factor is not same as in the Perkins&Hage.
Originally posted by Crumpp
Those calculations are confirmed by another professional aeronautical engineering firm that is producing FW-190's.
At least Mr. Colling from the Flugwerk told me that he has not given out any comments on e factor.
Originally posted by Crumpp
In the other thread your willingness to manipulate data on the Spitfire was very evident.
Ah, now you want to talk about other threads again, this thread is about determining e factor. All I can see here is that you continously change your story, fail to prove your arguments, give contradictory statements and create od theories on lift coefficient (see above). You have continously done similar things in other threads too.
Generally I have pointed source for all data I have used in this thread and I have given the calculations as well. You have not.
Originally posted by Crumpp
So tell me what speed is that value for "K" calculated under?
Actually the K values in the FW paper are for certain flying condition, not for exact speed. The report says climb speed ie around 250-300km/h TAS near sea level in the case of the full scale plane. The High speed value is roughly for top speed.
Originally posted by Crumpp
What speed is the spitfires?
Get the reports, the sources are given above. Generally I have used below mach 0,5 values for all calculations ie these should be quite comparable if wanted. Overall the exact speed (if below Mach 0,5) is not important if we know the Cl and/or value of the K.
gripen
-
So now you change your story again. Above you claim that e factor averages 0,87 which is actually nonsense because value of the K gives directly exact value of the e, there is no need to calculate average or fit the line as alternative method.
Here you drop some complete falsehood. My story remains the same.
At least Mr. Colling from the Flugwerk told me that he has not given out any comments on e factor.
My God man! Are you running around bugging folks over this? Mr Colling and I corresponded some months ago. His answer was he would be glad to help me out but to contact him during a specific time frame later in the year. That time frame has not elapsed. At this point his data will simply confirm what I already have on hand.
Now, It is NONE of your business at this point who I am contacting. You want to find out my sources, then buy the book or DON"T for that matter.
Gripen, I just can't throw everything I find out on the table because it would leave nothing for the book
Who actually did the calculation and how?
See above.
or the definition of the e factor is not same as in the Perkins&Hage.
The numbers and formulas can produce variations. That is exactly what the NASA engineer warned you about. If you don't know the formulas behind the values produced you cannot begin to interchange values into different formulas. They are not interchangeable. The scales are different.
Not only that but the Polar Plots change as well. I have at least three different sets of polar plots (Rechlin and Focke-Wulf) on the FW-190. All are somewhat different.
That is exactly why I am hiring professionals and using only primary sources.
Using the same scale, the FW-190 is only a fraction behind the Spitfire in wingtip efficiency.
Generally I have pointed source for all data I have used in this thread and I have given the calculations as well. You have not.
No you haven't. In the other thread you blatantly substituted data for the Spitfire Mk VIII and claimed it to be a Mk IX. Throughout my dealings with you are more than willing to manipulate data to present your point of facts instead of allowing the facts to define your world.
The report says climb speed ie around 250-300km/h TAS near sea level in the case of the full scale plane. The High speed value is roughly for top speed.
Exactly. A 50 Km window at climb speed and a "roughly" for top speed at what altitude? Hardly exact and hardly worthy of comparison.
Overall the exact speed (if below Mach 0,5) is not important if we know the Cl and/or value of the K.
If you want to compare planes it certainly is important. Nobody cares that when the FW-190 and the Spitfire have the same CL the spitfire's wing is more efficient. What matters is under the SAME conditions of flight.
Just like that NASA engineer told you.
Crumpp
-
Originally posted by Crumpp
Here you drop some complete falsehood. My story remains the same.
Let's look at the evolution of your story:
Crumpp 9-21-2004 11:58 AM:
"Just heard back from Flugwerk. They are building FW-190A8/N's and have restored several. According to them the "e" factor from FW documentation is:"
Crumpp 10-28-2004 02:40 PM:
"Those calculations are confirmed by another professional aeronautical engineering firm that is producing FW-190's."
Crumpp 10-30-2004 02:48 PM:
"Mr Colling and I corresponded some months ago. His answer was he would be glad to help me out but to contact him during a specific time frame later in the year. That time frame has not elapsed. At this point his data will simply confirm what I already have on hand."
Basicly you drop some complete falsehood.
Originally posted by Crumpp
My God man! Are you running around bugging folks over this?
Just look their site, they have actually put up a "rumour controll" because some people tend to put words in their mouth just like you did above.
Originally posted by Crumpp
You want to find out my sources, then buy the book or DON"T for that matter.
So why are you in this discussion if you can't prove your statements?
Originally posted by Crumpp
The numbers and formulas can produce variations.
Actually we talk about just one exact dataset which gives exact value of the K, there is no variaton.
Originally posted by Crumpp
In the other thread you blatantly substituted data for the Spitfire Mk VIII and claimed it to be a Mk IX.
Please point me where I have done such claim.
Originally posted by Crumpp
A 50 Km window at climb speed and a "roughly" for top speed at what altitude?
Well, you just can't understand the linear stage of the Cd/Cl^2 curve.
gripen
-
"Just heard back from Flugwerk. They are building FW-190A8/N's and have restored several. According to them the "e" factor from FW documentation is:"
No Your coming in the middle of my business and trying to claim you understand.
They did send me some documentation and only the results. They have not sent me the entire thing or how they arrived at that result.
Please point me where I have done such claim.
http://www.hitechcreations.com/forums/showthread.php?s=&postid=1391720#post1391720
So why are you in this discussion if you can't prove your statements
Again your implying a falsehood. Just because I don't post data to prove anything to you does not mean it is not available. Buy the book.
Just look their site, they have actually put up a "rumour controll" because some people tend to put words in their mouth just like you did above.
I highly doubt it is from the one email correspondance I have made with them. They are extremely nice and very helpful folks who immediately contacted me and provided an answer with a time frame to recontact them for details. I will contact them again during the timeframe they specified and not before.
Again, this is just a simple attempt to "smear".
Please point out your claim about thier site:
http://www.flugwerk.com/
Actually we talk about just one exact dataset which gives exact value of the K, there is no variaton.
Again totally not relevant to a comparision under the same flight conditions of the Spitfire and the FW-190.
Crumpp
-
Originally posted by Crumpp
No Your coming in the middle of my business and trying to claim you understand.
Plain nonsense.
Originally posted by Crumpp
http://www.hitechcreations.com/forums/showthread.php?s=&postid=1391720#post1391720
Funny, you linked your own post.
Originally posted by Crumpp
Please point out your claim about thier site:
http://www.flugwerk.com/
The Rumour control is in the diary page.
Originally posted by Crumpp
Again totally not relevant to a comparision under the same flight conditions of the Spitfire and the FW-190.
This has nothing to with the Spitfire. The FW datasheet gives exact value of the K and that results e factor 0,78 at linear stage of the Cd/CL^2 curve.
gripen
-
I lost the track. Can somebody fill me in, what was the original question? :aok
-
Just look their site, they have actually put up a "rumour controll" because some people tend to put words in their mouth just like you did above.
What are you talking about Gripen? Your throwing out complete lies now. The rumours are about Flugwerk's financial status and have nothing what so ever to do with anything in this or any GAME.
As for the material I requested from them. It is a business transaction as I am paying them for their services and experience with the type.
Plain nonsense.
Exactly what your pedaling.
Here Look closely at their website diary page:
http://www.flugwerk.com/new/diary/diary.shtm
Try and find anything remotely close to your nonsensical accusations.
Funny, you linked your own post.
And yours as well. It in the thread towards the end. Right before you panicked and started this one.
Crumpp
-
Originally posted by Crumpp
What are you talking about Gripen? Your throwing out complete lies now.
You claim above that the quote is from Flugwerk and that your numbers are confirmed by them. These are not true.
Originally posted by Crumpp
And yours as well. It in the thread towards the end. Right before you panicked and started this one.
Please show me the post where I "blatantly substituted data for the Spitfire Mk VIII and claimed it to be a Mk IX" as you claimed above.
gripen
-
Gripen says:
You claim above that the quote is from Flugwerk and that your numbers are confirmed by them. These are not true.
Email Crumpp recieved says:
I was checking my documents for the FW190 and the tip efficiency factor, or rather the Oswald Ellipse factor is about 1.15 or 0.87 depending on how you look at it. Basically it has about 15% more induced drag than a theoretical elliptical lift distribution.
Again Gripen,
You have to compare the A/C under the same flight condition. The "e" factor will change. Your comparison is not valid since you cannot reproduce the "e" factor for the Spitfire under the same conditions.
Crumpp
-
Well, Crumpp's quote is not from the Flugwerk as he self actually admits above.
Originally posted by Crumpp
You have to compare the A/C under the same flight condition. The "e" factor will change. Your comparison is not valid since you cannot reproduce the "e" factor for the Spitfire under the same conditions.
This has nothing to with comparisons or exact flight conditions, the FW datasheet gives value of 1,24 for K at linear stage of the Cd/Cl2 curve. And that results e factor 0,78 at linear stage of the Cd/CL^2 curve. There is no variation nor other answer.
gripen
-
This has nothing to with comparisons or exact flight conditions, the FW datasheet gives value of 1,24 for K at linear stage of the Cd/Cl2 curve. And that results e factor 0,78 at linear stage of the Cd/CL^2 curve. There is no variation nor other answer.
It has everything to do with exact flight conditions, Gripen.
There are portions of the flight envelope were the total drag of the Spitfire is greater than the FW-190's and vice versa. The "e" factor is in fact listed on the document farther down the column.
Crumpp
-
Originally posted by Crumpp
It has everything to do with exact flight conditions, Gripen.
As noted several times above, the exact Cl value or speed value are not so important if the value of the K is for linear stage of the Cd/CL^2 curve and the speed is below compressibility speeds (say less than mach 0,5).
(http://personal.inet.fi/koti/soon.moro/ERREE.JPG)
As can be seen and read from this scan from Perkins&Hage, the slope of the Cd/CL^2 curve is about linear between the CL^2 values 0,2 and 1,0. At climb speeds (say 250-300 km/h) the CL^2 value of the Fw 190A is around 0,25-0,5 so we don't need to know exact speed and we also know that the speed is not at compressibility area. Same is true for all the wind tunnel data I have used for my calculations.
This thread is not about comparing planes but about determining the values of the e factor of WWII fighters. Based on my calculations I can say with good certainty that the generic aspect ratio based formulas seem to give about 10% higher values of e factor than wind tunnel data. Flight tests seem to give even smaller values of e but I have too little data to say more, possibly most of the turn tests were done very close to Clmax ie not in the linear stage of the Cd/CL^2 curve.
Originally posted by Crumpp
The "e" factor is in fact listed on the document farther down the column.
Hm... at least my copy just lists detailed drag data for components, flying surface data and propeller and thrust data. Apparently you have just the headlines as you linked above.
Besides, as can be seen from the above image from the Perkins&Hage, the value of the K is all we need (which I have linked above), it gives directly slope of the Cd/CL^2 curve because:
K = 1 / (pi * AR * e)
gripen
-
Hm... at least my copy just lists detailed drag data for components, flying surface data and propeller and thrust data. Apparently you have just the headlines as you linked above.
Nope. I have four other pages that come with it entitled "Polarenmessung in Chalais-Meudon".
Crumpp
-
Originally posted by Crumpp
Nope. I have four other pages that come with it entitled "Polarenmessung in Chalais-Meudon".
Well, that's another report. You are most wellcome to bring in data in if you really want to discuss about the determining of the efficiency factor. The problem is that you have claimed many kind of things in this thread without anykind of proof.
gripen
-
The problem is that you have claimed many kind of things in this thread without anykind of proof.
With out any claim? Sorry Gripen, the guys who did those calculations are working aeronautical engineers. Now I don't have the break down on how they arrived at that yet but when the time is right I will.
I am paying for some aeronautical engineering firms to conduct an analysis off the data as it is impartial and much more accurate.
I don't want to discuss anything with you Gripen. You manipulate data and have no credibility.
http://www.hitechcreations.com/forums/showthread.php?s=&threadid=128143&perpage=50&pagenumber=4
Crumpp
-
Originally posted by Crumpp
I don't want to discuss anything with you Gripen.
So why in earth you are in this thread?
This thread is about determining efficiency factor, not about your supposed sources, nor about your supposed book nor about someone's credibility.
Originally posted by Crumpp
You manipulate data and have no credibility.
Could you point where exactly I manipulate data and how? At least I could not find anything such from your linked page.
Regarding credibility, just look at your own posts in this thread; great logic overall.
Originally posted by Crumpp
I am paying for some aeronautical engineering firms to conduct an analysis off the data as it is impartial and much more accurate.
Well, If someone is getting something else from the above linked values of the K, you are wasting money.
gripen
-
Emmm.
"--------------------------------------------------------------------------------
Originally posted by Crumpp
You manipulate data and have no credibility.
--------------------------------------------------------------------------------"
Just ran over this thread, but can't spot that.
Isn't it getting a tad too hot?
-
Angus,
You have to look at the evolution of Gripen's flat plate calculations in the thread I linked too. If you don't do the math you won't see it. I don't make those accusations lightly.
Compare his values to the values calculated for ALL the Spitfire Mk IX's for flat plate area. NONE of the Mk IX's have less flat plate area than the FW-190. That was when I realized what was going on.
However the Spitfire Mk VIII does. If you look at Gripens calculations then it becomes clear. That is the main reason I have no interest in continuing any kind discussion with him on this or any subject.
Gripen is a smart guy and has gotten by because many folks don't understand what he is doing.
However, It's not productive nor factual to continue this. It's data manipulated for advantage in a game. It has nothing to do with reality or history.
HTC is being provided with copies of the original reports and can make their own conclusions.
someone's credibility.
It's all about credibility. You don't have any in my book, Gripen.
I'm sure you will try and throw up some fluff making claims I don't understand your calculations. We both know the truth though.
Crumpp
-
Whoops.
I am not an aerodynamic expert, nor have I dipped into this exact thread so much, but it seems as if it may be worth it.
BTW, (silly question, but....gives me a quicker start) How would you best define flat plate area?
The Spitfire VIII and IX are very very similar, with the VIII having a cleaner line (little odds and ends like retractable tailwheel), but no big difference.
At the back of my head I am bothered by the memory of those having a slightly different wing, but I haven't been able to dig this up in my books yet. There could possibly be some difference here?
(I read too much and take too few notes, - well it's a good gymnastic for the brain)
Anyway, letts have a closer look.
-
Originally posted by Crumpp
Compare his values to the values calculated for ALL the Spitfire Mk IX's for flat plate area. NONE of the Mk IX's have less flat plate area than the FW-190.
Well, this has nothing to do with subject of this thread so I moved this back to the original thread.
Originally posted by Crumpp
It's all about credibility. You don't have any in my book, Gripen.
I'm sure you will try and throw up some fluff making claims I don't understand your calculations. We both know the truth though.
All I see here is that you try to talk about a subject that you don't understand at all. Basicly you are continously throwing in accusations and various claims without anykind of proof.
gripen
-
I think your getting your values confused and guessing at the German's notations.
Otherwise the FW-190A8 thru FW-190D12 all have the exact same value of K.
Just doesn't hold out when you have the polar plots. The values are different.
That is why I am having the entire windtunnel test report professionally translated. Not because the data on the FW-190V5g will be any good but because it goes into detail on the formulas the Focke-Wulf Design team used.
That tops off the fact you are taking the value of K from the ends of the envelope. One is the best climb speed and the other is top speed in level flight. Look up the definition of best climb speed. It is hardly on the linear portion. Any lower and it's in the stall. The opposite end of that spectrum is top level flight speed.
Now at the low speed end of the envelope the Spitfire will have the advantage in drag.
Crumpp
-
Originally posted by Crumpp
I think your getting your values confused and guessing at the German's notations.
Otherwise the FW-190A8 thru FW-190D12 all have the exact same value of K.
The wing is same and the drag values are different, that's quite logical indeed.
Originally posted by Crumpp
That tops off the fact you are taking the value of K from the ends of the envelope. One is the best climb speed and the other is top speed in level flight. Look up the definition of best climb speed. It is hardly on the linear portion. Any lower and it's in the stall. The opposite end of that spectrum is top level flight speed.
Total nonsense, the climb Cl at 250-300km/h is around 0,5-0,7 depending on conditions and that's 0,25-0,49 in Cl^2 scale ie clearly in the linear stage. And far from the stall.
gripen
-
Total nonsense, the climb Cl at 250-300km/h is around 0,5-0,7 depending on conditions and that's 0,25-0,49 in Cl^2 scale ie clearly in the linear stage. And far from the stall.
You need to learn the definition of best climb speed:
http://www.av8n.com/how/htm/energy.html#fig-power-curve-regimes
The dividing line between the mushing regime and the front side of the power curve is the highest point on the power curve. At this point, the airplane can fly with the minimal amount of dissipation; this is the “low-rent district”. The airspeed where this occurs is called the best-rate-of-climb airspeed and denoted
It is not in the linear portion of the curve which is:
The right-hand part of the curve (from moderate airspeeds on up) is called the front side of the power curve. Normal cruising flight is conducted in this range of airspeeds.
The front side of the power curve.
Crumpp
-
Originally posted by Crumpp
You need to learn the definition of best climb speed....
Well, now this is going to be fun, your link has nothing to do with the drag polars nor with the linear stage of the Cd/Cl^2 curve.
The Cl for the climb speed is easy to calculate from the climb speed and theangle of the climb and it does not need to be accurate if it's in the linear stage of the Cd/Cl^2 curve.
If you want to argue, you should prove that the Cl is less than say 0,2-0,3 or more than say 1 at best climb speed; between these values the Cd/Cl^2 curve is about linear.
gripen
-
Well, now this is going to be fun, your link has nothing to do with the drag polars nor with the linear stage of the Cd/Cl^2 curve.
Gripen,
My contention is that without the drag polars you are just guessing.
Crumpp
-
One little question.
Did you guys ever fly a single engined monoplane yourselves, climbing, banking and so on?
Just curious........
-
One little question.
Yep.
Check your email Angus.
Crumpp
-
Originally posted by Crumpp
My contention is that without the drag polars you are just guessing.
Nonsense, the formula:
CwF = Cw0 + K * Ca^2
is directly the drag polar.
Basicly you just continously throw in nonsense arguments.
gripen
-
Basicly you just continously throw in nonsense arguments.
What is ridiculus in saying that without the polars you are just guessing?
Besides I will take a trained aeronautical engineer any day over your word especially after seeing your refusal to use Luftwaffe tested data in your calculations.
As long as you:
1. Insist on using the taper ratio of a theoritical ellipse for the Spitfire when every aeronautical engineering student knows the wing twist destroyed the elliptical wet lifting surface of the wing.
2. Do use the drag polars.
Your values will be wrong. Garbage in, Garbage out.
Crumpp
-
Originally posted by Crumpp
What is ridiculus in saying that without the polars you are just guessing?
The problem is that you can't under stand that the formula:
CwF = Cw0 + K * Ca^2
is directly drag polar.
Originally posted by Crumpp
1. Insist on using the taper ratio of a theoritical ellipse for the Spitfire when every aeronautical engineering student knows the wing twist destroyed the elliptical wet lifting surface of the wing.
As noted several times in these discussions, you are the only one who has claimed that the Spitfire has e factor value 1. In addition you have quoted couple contradictory claims for near elliptical lift distribution taper ratio values (0,4 and 0,52 and actually 0,67 for elliptical wing ;))
Originally posted by Crumpp
2. Do use the drag polars.
Well, it seems that I'm the only one in this thread who has actually calculated e factors from the drag polars.
gripen
-
Well, it seems that I'm the only one in this thread who has actually calculated e factors from the drag polars.
Yep. For the Spitfire, Me-109, and P51. Unfortunately you don't have the FW-190 drag polars and your not getting them from me.
In addition you have quoted couple contradictory claims for near elliptical lift distribution taper ratio values (0,4 and 0,52 and actually 0,67 for elliptical wing )
Exactly. Why is that? Because the taper ratio for the Spitfire is impossible to nail down since the wet lifting area in no way matches the elliptical shape of the wing due to the wing twist destroying the benefits of the elliptical tips. That leaves a wide margin that delivers correct values that are meaningless in the real world. Very open to data manipulation.
On the other hand, being a much later design and it's designers being well aware of the benefits of elliptical distribution purposely twisted the FW-190's wing to achieve it.
In the real world they were probably pretty close as the 1965 formulas compute them.
In other words without tested drag polars under the same conditions it's impossible to correctly compare performance.
A general formula that does not include taper ratio in this case, as Badboy pointed out, probably yields results that are much closer to real world performance.
Crumpp
-
Originally posted by Crumpp
Unfortunately you don't have the FW-190 drag polars and your not getting them from me.
Well, the Fw document gives the value of the K and that's all I need. But maybe I'll dug that "Chalais-Meudon" report from somewhere to have even more fun.
Originally posted by Crumpp
Because the taper ratio for the Spitfire is impossible to nail down since the wet lifting area in no way matches the elliptical shape of the wing due to the wing twist destorying it. That leaves a wide margin that delivers correct values that are meaningless in the real world. Very open to data manipulation.
Plain nonsense as usual, so now you introduce "the wet lifting area" ;)
Originally posted by Crumpp
In other words without tested drag polars under the same conditions it's impossible to correctly compare performance.
So here you admit again that you don't understand at all the linear stage of the Cd/Cl^2 curve.
Originally posted by Crumpp
A general formula that does not include taper ratio in this case, as Badboy pointed out, probably yields results that are much closer to real world performance.
Well, the generic formulas seem to assume tapered wing so these should work in those cases but as pointed out above, those formulas seem to give too high values even in those cases. If this kind of formula gives about correct result for other kind of wing, it's more or less luck.
gripen
-
Crumpp, are you referring to the twist as washout?
-
Originally posted by Angus
I'd say NO.
Aspect ratio is only a function of chord and span.
You'd be missing shape, wingtip shape, thickness and dihedral.
I agree.
-
Well, the generic formulas seem to assume tapered wing so these should work in those cases but as pointed out above, those formulas seem to give too high values even in those cases. If this kind of formula gives about correct result for other kind of wing, it's more or less luck.
As opposed to just guessing at the correct taper ratio? Your stand is that is more accurate?
Please.
In addition you have quoted couple contradictory claims for near elliptical lift distribution taper ratio values (0,4 and 0,52 and actually 0,67 for elliptical wing )
Which still leaves this unanswered. What is the guess for the correct taper ratio of the Spitfire?
Crumpp
-
What Spitfire?
I presume the most produced wing, i.e. from Mk I through MkV all the way to many MK IX's right?
Anyway, you should be able to calculate the taper ratio from chord and span and for a perfect ellipse, which however the Spitfire did not have, but very very close.
-
Originally posted by Crumpp
As opposed to just guessing at the correct taper ratio?
Nonsense, there is no need to use generic formulas when the value of the K is known.
Originally posted by Crumpp
Which still leaves this unanswered. What is the guess for the correct taper ratio of the Spitfire?
Nonsense, there is no need to "quess" when the shape of the drag polar is known (ie the value of the K).
gripen
-
Nonsense, there is no need to use generic formulas when the value of the K is known.
Not talking about the FW-190's. I am refering to the Spitfires. Nice bait and switch though.
Nonsense, there is no need to "quess" when the shape of the drag polar is known (ie the value of the K).
Again I was refering to the taper ratio of the Spitfire.
The one you keep asking me to decide on here:
In addition you have quoted couple contradictory claims for near elliptical lift distribution taper ratio values (0,4 and 0,52 and actually 0,67 for elliptical wing )
So which "guess" should I take?
Anyway, you should be able to calculate the taper ratio from chord and span and for a perfect ellipse, which however the Spitfire did not have, but very very close.
Actually according to Lednicer it was not very close. It was slightly better than the other two but not a fantastic lead. Certainly nowhere near a perfect ellipse.
Crumpp
-
Originally posted by Crumpp
Not talking about the FW-190's. I am refering to the Spitfires. Nice bait and switch though.
Nonsense, wind tunnel data of the Spitfire gives the shape of the drag polar. And the Fw data gives the value of the K. There is no need to quess.
Originally posted by Crumpp
The one you keep asking me to decide on here:
Nonsense, you have quoted two different values for taper ratio which gives near elliptical lift distribution. It's your problem.
Originally posted by Crumpp
So which "guess" should I take?
You made those contradictory statements, it's your own problem.
Originally posted by Crumpp
Actually according to Lednicer it was not very close. It was slightly better than the other two but not a fantastic lead.
Nonsense, Lednicer analyzed span loading of the wing but he did not give any values.
gripen
-
Nonsense, you have quoted two different values for taper ratio which gives near elliptical lift distribution. It's your problem.
Nonsense,
Your Spitfire calculations are Nonsense because you use the taper ratio for a theoretical ellipse when the Spitfire does not have a perfect elliptical wingtip. The twist ruined it.
Nonsense, you have quoted two different values for taper ratio which gives near elliptical lift distribution. It's your problem.
Nonsense, I quoted a range for values.
You made those contradictory statements, it's your own problem.
Nonsense, Your English is not good enough to understand what I wrote. Get a better dictionary.
Nonsense, Lednicer analyzed span loading of the wing but he did not give any values.
Learn to read again. Lednicer commented on the efficiency factor. The Spitfire is only "slightly ahead" and ruined the benefits of elliptical construction by twisting the wing to fix the harsh stall.
Crumpp
-
Originally posted by Crumpp
Your Spitfire calculations are Nonsense because you use the taper ratio for a theoretical ellipse when the Spitfire does not have a perfect elliptical wingtip. The twist ruined it.
If I know the shape of the drag polar, the value of the K can be easily determined as pointed out in the PerkinsHage. There is no need to quess or even know the taper ratio.
Shortly you have no idea what you are talking about.
Originally posted by Crumpp
I quoted a range for values.
No, you quoted two contradictory claims for taper ratios which should give near elliptical lift distribution:
Your post 09-11-2004 02:50 PM:
"However, the formulae presented by Professor Wood is for a taper ratio of 0.57 which yields an almost elliptical lift distribution, "
Your post 09-18-2004 04:01 AM:
"A taper ratio of 0.4 most closely approximates the elliptical."
Originally posted by Crumpp
Learn to read again. Lednicer commented on the efficiency factor.
Here is the direct quote again:
"loading distribution is not elliptical, though it is probably the most optimum of the three from the induced drag standpoint".
Besides, you are the only one in these discussions to claim e factor 1 for the Spitfire as can be verified from here (http://www.hitechcreations.com/forums/showthread.php?s=&postid=1368542#post1368542).
gripen
-
Of this:
"Actually according to Lednicer it was not very close. It was slightly better than the other two but not a fantastic lead. Certainly nowhere near a perfect ellipse. "
Building a perfect ellipse is easily doable, but to get there, the aspect ratio will suffer instead of the induced drag gain by getting all the way to the ellipse.
Why do you think Mitchell picked this exact design?
-
Building a perfect ellipse is easily doable, but to get there, the aspect ratio will suffer instead of the induced drag gain by getting all the way to the ellipse.
Read Lednicer's quote on the fact designers were well aware of the benefits of elliptical distribution and it was easily achieved by twisting the wings. This occurred in the later designs of the FW-190 and the P51.
The Spitfire on the other hand had an elliptical distribution already. That distribution caused a nasty and dangerous stall. To change the stall characteristics so they were workable, the Spitfire designers twisted the wing to cut down on the elliptical distribution.
Thereby bringing all three aircraft in a similar catagory for wingtip efficiency. Sure the Spitfire has the lead but it's not a commanding lead for this characteristic. It's lower wingloading is the main contributor to it's fighting characteristics not it's efficiency factor.
Crumpp
-
I don't have that report, but I'd like too. :):):):)
Anyway, from some source (think it was a pilot) I heard that actually the Spitfire wing would NOT stall all at once, which would then be a normal sideffect, since a complete ellipse should stall all at the same time. Am I right there?
The genius part with the Spitfire is perhaps exactly there, it's a clever compromize to acquire low wing loading with the lowest possible drag. The low wingloading pays for itself instead with good climb and good turnrate and good ceiling.
The second genius part would than be that it's almost uncompressable, not sure why exactly though.
The penalty: Less absolute top speed.
-
Anyway, from some source (think it was a pilot) I heard that actually the Spitfire wing would NOT stall all at once, which would then be a normal sideffect, since a complete ellipse should stall all at the same time. Am I right there?
That is exactly why it was twisted. To destroy some of the ellipse.
Check your email.
Crumpp
-
Originally posted by Crumpp
Read Lednicer's quote on the fact designers were well aware of the benefits of elliptical distribution and it was easily achieved by twisting the wings. This occurred in the later designs of the FW-190 and the P51.
While some amount of washout can improve lift distribution of the rectangular wing, it can be seen from Lednicer's lift distribution chart that the Fw 190 and P-51 have too much washout for optimal lift distribution:
(http://personal.inet.fi/koti/soon.moro/sp.jpg)
With less washout the lift distribution would have been much closer elliptical. Therefore it's easy to see that reason for the amount of wing twist in the Fw 190 and P-51 is same as in the Spitfire; to get the wing root stall before the tip.
Originally posted by Crumpp
Thereby bringing all three aircraft in a similar catagory for wingtip efficiency.
The lift distribution chart is quite clear, the Spitfire has "the most optimum of the three". This can be also verified from the wind tunnel and FW data.
Originally posted by Crumpp
It's lower wingloading is the main contributor to it's fighting characteristics not it's efficiency factor.
Well, here I agree. If we compare planes, generally the one with lowest wing loading has lowest drag rise when the g load increases. But that is a bit out of subject of this thread.
gripen
-
Thanks for clipping the chart from the article. It illustrates perfectly what is going on with the FW-190 vs Spitfire wing.
At low speeds the Spitfire does have the advantage. At high speeds the efficiency gap closes dramatically. Just as the Spitfire moves farther away at high speed from the "elliptical shape", the FW-190 moves closer and eventually surpasses the Spitfire in the last bit of the envelope.
According to your calculations that never happens. According to Lednicers data, the drag polars, and other aeronautical engineers figures it does.
(http://www.onpoi.net/ah/pics/users/503_1100203303_span2.jpg)
Crumpp
-
Originally posted by Crumpp
Thanks for clipping the chart from the article. It illustrates perfectly what is going on with the FW-190 vs Spitfire wing.
Well, I don't know why you want to make yourself laughable but "´Semispan Fraction" means relative location along the wing half. It has absolute nothing to do with the speed.
gripen
-
Hi,
Here is an explanation which may or may not have the level of detail you require. Lift is generally described mathematically, or through the mathematical abstraction, as circulation. This circulation is a function of the shape of the wing, not the wing area vs span ratio. As an example of how the wing Coefficient of lift varies along its span, here is a url: http://www.nurflugel.com/Nurflugel/Horten_Nurflugels/theory/theory.html
The concept of vorticity and circulation is not physically easy to grasp, and I make no claim to have that insight. However, circulation is what creates the lift in the inviscid (non viscous) flow regimes. This circulation theory provides an elliptical distribution (for vorticity) giving a minimum resistance due to vortex downwash. This elliptical distribution is for the wing along its span and the aspect ratio can vary.
My apologies if all this is way too abstract but it is how I was taught. There are several papers by Prandtl from 1918 to the early 20's that are on the NACA server at (http://naca.larc.nasa.gov). He first made this connection in mathematical terms between wing shape, lift, and drag.
Regards,
Malta
-
Thanks Stantond. Great explaination. Now to apply it to this chart.
So one end is the wingtip, correct? I certainly don't see where the Spitfire has any advantages if that is the case.
Crumpp
-
Originally posted by Crumpp
So one end is the wingtip, correct? I certainly don't see where the Spitfire has any advantages if that is the case.
At least Lednicer, me and probably everyone else can see the diffrerence the Spitfire "being the most optimum of the three".
gripen
-
How can twisting an elliptical wing destroy it any more than twisting a tapered wing?
After all, that part is about washout right? And that is a drag issue more than a lift issue.
I always thought the Spitfire wing was not a complete ellipse in shape, i.e. a total ellipse would have meant a bit wider chord.
Or am I in the dark here?
-
Originally posted by Angus
How can twisting an elliptical wing destroy it any more than twisting a tapered wing?
It's quite simple; without washout elliptical wing has optimal lift distribution and washout makes it less optimal. Lift distribution of the tapered wing can be improved with some amount of washout (in some cases) but in the case of the Fw 190 and the P-51 it can be clearly seen that that these have too much washout for optimal lift distribution.
Basicly in the all three cases the main reason for the washout is to improve stalling characters of the plane and in all three cases amount of washout decrease efficiency of the wing.
gripen
-
So appearantly the best way is to use wings with slats and no washout to have both optimum stall characterstics AND good lift effiency. ;)
-
Originally posted by Kurfürst
So appearantly the best way is to use wings with slats and no washout to have both optimum stall characterstics AND good lift effiency. ;)
That's one way to do it. The e factor of the Bf 109 appear to be in the same ballpark with the P-51 and Fw 190 but it generally had the better stall characters.
gripen
-
And then enters the dihedral :D
-
Originally posted by Angus
And then enters the dihedral :D
Hmm, the dihedral (V-stellung) supposed to increase the stability of the aircraft, this being important for low mounted wings, ie. fighters with neutral roll axis stability... is there anything else?
-
Depending on which end is the wingtip....
(http://www.onpoi.net/ah/pics/users/503_1100271849_span2.jpg)
The spitfire does have a more efficient wing on the whole that was never in doubt. It is the wingtip though that determines induced drag though and the twist destroyed the elliptical shape of the lifting surface at the tip. You can see that by this chart.
Just as David Lednicer says. The Spitfire is the most optimum of the three. It does not have the commanding lead you want to give it, nor does always have less drag than the 190 no matter what portion of the flight envelope. As you picture it.
And then enters the dihedral
The Spitfires roll was not as stable??
Crumpp
-
In fact I would veture to say that if this guy:
http://www.grc.nasa.gov/WWW/K-12/airplane/benson.html
Doesn't know a way to calculate wingtip efficiency or gain any insights into it off of the above chart. I would say we are spinning our wheels with this.
Gripen didn't you use this chart to figure out the FW-190's wingtip efficiency earlier in this thread?
My question:
Is there anyway to figure out the exact Wingtip efficiency factor of the Aircraft using this chart?
Not asking you to figure it out. Just confirm whether you could do it off this chart or gain any insight. Using the forumla for Oswalds Efficiency factor, the FW-190, P51D and the Spitfire are extremely close.
Tom says:
There probably is .. but I don't know it.
Crumpp
-
Well, one can draw lines to the picture but the spanwise efficiency is function of the whole wing despite it is called some times wing tip efficiency. Besides, we talk here about efficiency of the whole airframe as in Perkins&Hage.
Regarding efficiency calculation based on relative lift distribution, it certainly gives relative difference of lift distribution but it takes account only the differences in induced drag, not the viscous. Still, for one reason or another it seem to give suprisingly similar results as polar analysis.
gripen
-
Still, for one reason or another it seem to give suprisingly similar results as polar analysis.
Yeah I am sure we could throw some numbers out there that could be remarkably similar to polar analysis. Does not mean they are correct by any means.
What is that saying, an infinite number of Monkeys banging away on an infinite number of typewriters will eventually reproduce the works of Shakespeare...
Crumpp
-
You can easily test it yourself, besides FW data gives the drag polar:
CwF = Cw0 + K * Ca^2
No need to quess anything.
gripen
-
You should check the polars.
Crumpp
-
You are most wellcome to post polars here if those are somehow different than given in FW data.
gripen
-
You are most wellcome to post polars here if those are somehow different than given in FW data.
No, Thanks. I sent them in to Pyro. Positive HTC will do a fair and impartial set of calcs off them.
Crumpp
-
"Why do you think Mitchell picked this exact design?"
Mitchell picked that design because he wanted as large wing area as possible and for certain reasons it is the most sensible choice there is if the wing area is considered the most important factor.
You can get nearly the same Cl efficiency from a rectangular wing as seen in many modern day small a/c but the elliptic wing is better for a fighter aircraft because you can build the tip of the wing more rigid than it would be possible in rectangular shape plus the elliptic shape produces less drag.
The negative effect of the elliptic shape is the bad stability of the design in other than level flight which was probably why the Spit had such tall and slim fuselage to act as a huge rudder to counter that imbalance. Further maneuvering stability was produced by the washout which made the wing act more like a rectangular wing as seen in P51 or FW without sacrificing much of the lift qualities of elliptic design as seen in figure presented earlier in this thread.
It should also be noted that that figure shows the Cl for level flight AoA so as AoA increases the Cl for the wingtip rises to , more or less, meet the Cl values of the optimum elliptic profile while the Cl for the root may have already lost its lift due to stall.
-C+
-
Emm...Charge....wooot?
"You can get nearly the same Cl efficiency from a rectangular wing as seen in many modern day small a/c"
Well, with 30% more induced drag for a full rectangular.
"The negative effect of the elliptic shape is the bad stability of the design in other than level flight"
Never heard that one.
The Spitty had C of G problems, but they had nothing to do with the ellipse.
And:
"Further maneuvering stability was produced by the washout which made the wing act more like a rectangular wing as seen in P51 or FW without sacrificing much of the lift qualities of elliptic design as seen in figure presented earlier in this thread. "
The 190 and P51 have tapered wings, not rectangular. Tapered wing is the cheap way of obtaining something close to the elliptical effect.
Finally, I think I've heard before that the Spitfire wing root stalled before the tips. Anybody?
-
Seems that Mitchell ended to the elliptic form for several reasons and need to place the wheels, guns and ammunition inside wing are probably among these. Otherwise, it's quite probably that they did not think mass production at all; in the mid thirties very few types reached large production numbers.
Regarding stall, the Spitfire had generally good stall characters like the Bf 109 ie no tendency to tip stall.
gripen
-
Regarding stall, the Spitfire had generally good stall characters like the Bf 109 ie no tendency to tip stall.
Yes it did. AS Lednicer points out however it would have had a very dangerous stall without the twist.
Although it sacrificed some wingtip efficiency for better stall characteristics, I think the tradeoff was worth it.
It allowed poorly trained pilots to survive and gain experience quickly. As the NACA trials point out, the Spitfire exhibited an amazing amount of control and stability even during the stall.
Crumpp
-
Ummm?
"--------------------------------------------------------------------------------
Regarding stall, the Spitfire had generally good stall characters like the Bf 109 ie no tendency to tip stall.
--------------------------------------------------------------------------------
Yes it did. AS Lednicer points out however it would have had a very dangerous stall without the twist. "
Crumpp, your text needs a wee refining.
It was twisted a bit and not completely elliptical.
A wee more induced drag pro square, however maybe less square?, - anyway, the Spit stalled gently when compared to other WW2 planes of similar power.
The Hurricane would dip a wing BTW.
-
Your correct Angus. I was reading about the Spitfires development. Seems it was a windtunnel test wooden model in 1935 I was thinking of...:o
As far as the wingtip efficiency goes:
(http://www.onpoi.net/ah/pics/users/503_1100478548_spitfireelliptical.jpg)
Key word is probably.
Looking at the shape of the high pressure portion (red) of the wet lifting surface:
(http://www.onpoi.net/ah/pics/users/503_1100478910_spitwing.jpg)
The sharp ending is hardly elliptical. In fact the the FW-190's:
(http://www.onpoi.net/ah/pics/users/503_1100478959_fw190wing.jpg)
Looks alot closer (red). Taking into the lower pressure portion of the lifting surface (yellow), it sure looks to me like in the end they will come out very close.
Crumpp
-
Crumpp,
You have a very strong tendency to do yourself laughable, pressure distribution is not the same thing as lift distribution.
And there is no such thing as "wet lifting surface" except in your creative mind.
gripen
-
Emm...Charge....wooot?
->Sorry Angus, you obviously knew these things already but there are other people reading this thread, too?
"You can get nearly the same Cl efficiency from a rectangular wing as seen in many modern day small a/c"
Well, with 30% more induced drag for a full rectangular.
->Well, I said that elliptic produces less drag. 30%? Interesting.
"The negative effect of the elliptic shape is the bad stability of the design in other than level flight"
Never heard that one.
The Spitty had C of G problems, but they had nothing to do with the ellipse.
-> I think some of it may have. U gotta read more about elliptic wings maybe?
I believe it was a known profile even before Mitchell decided to use it but others chose not to, and there is a reason why. One thing is its complexity when considering mass manufacture but that is not the only reason.
Anyway Mitchell did a good job as history shows. The wing design was changed from Spitfire model 21 onwards (IIRC) for some reason. The later design probably rolled better because the aileron design was better in that one.
I forgot to mention that too much lift in the tip portion of the wing is probably not good if the torsional rigidity of the wing is not good enough as a flexing wing can have a negative effect on roll rate and stall characteristics in high speed or G loads. The wing needs to be torsionally stiff.
-C+
-
"pressure distribution is not the same thing as lift distribution"
If a wing creates lift due altering pressure around it so how does not pressure distribution have nothing to do with lift distribution?
Of course it depends on the vector of the pressure whether or not it produces lift...
If Spitfire had no washout I'd imagine that the red pressure area would resemble that of FW 190 and in higher AoA the red area will extend more to the tip of the wing. I guess the washout was taken into account in the model?
-C+
Edit: http://www.grc.nasa.gov/WWW/K-12/airplane/cp.html ??
-
It's quite simple, pressure distribution of the one side of wing tells absolute nothing about the lift distribution.
Basicly the pressure distribution tells how the pressure caused by the airflow varies around the airframe.
And yes, the model included washout.
gripen
-
It's quite simple, pressure distribution of the one side of wing tells absolute nothing about the lift distribution.
Where do you get that? Since induced drag is cause by an interaction of the pressure differentials between the TOP and the bottom of the wing it is safe to say that the Spitfire did NOT have an elliptical distribution by any means.
For a lifting wing, the air pressure on the top of the wing is lower than the pressure below the wing. Near the tips of the wing, the air is free to move from the region of high pressure into the region of low pressure. The resulting flow is shown on the figure by the two circular blue lines with the arrowheads showing the flow direction. As the aircraft moves to the lower left, a pair of counter-rotating vortices are formed at the wing tips. The line of the center of the vortices are shown as blue vortex lines leading from the wing tips. If the atmosphere has very high humidity, you can sometimes see the vortex lines on an airliner during landing as long thin "clouds" leaving the wing tips. The wing tip vortices produce a down wash of air behind the wing which is very strong near the wing tips and decreases toward the wing root. The local angle of attack of the wing is increased by the induced flow of the down wash, giving an additional, downstream-facing, component to the aerodynamic force acting over the entire wing. This additional force is called induced drag because it faces downstream and has been "induced" by the action of the tip vortices. It is also called "drag due to lift" because it only occurs on finite, lifting wings and varies with the square of the lift.
http://www.grc.nasa.gov/WWW/K-12/airplane/induced.html
No what is laughable is your continued insistance in the face of the facts that the Spitfire recieved some huge benefit from it's "elliptical" wing other than more room to put landing gear and of course more drag. At best it was barely ahead as the formulas show.
Crumpp
-
Well, wings it is.
The downwash btw pushes the tips upwards right?
The Hurricane had no dihedral, but I know of an instance where a pilot dived from 30K to the deck and got way to fast for a Hurricane.
He deformed wings, - they had a 4 degree dihedral afterwards, as well as wrinkles on the upper surfaces of the wing.
Looking for info on dihedral anyway.
Induced drag, is as you say, LIFT induced drag. All lift creates drag.
So the Spitfire's big wing both gives benefits and suffering.
It has a low wingloading which enables it to fly at lower A.o.A. at lower speeds, - A.o.A. creates drag.
Hence the maneuverability in the low speed region as well as good climb.
However, it has to pay at higher speeds.
-
Originally posted by Angus
The Hurricane had no dihedral,
The top of the wing might be flat but the lower side had dihedral.
How does this figure into every thing?
(http://mkmagazin.almanacwhf.ru/avia/hurricane_3.gif)
-
Not really sure.....
-
Induced drag, is as you say, LIFT induced drag. All lift creates drag.So the Spitfire's big wing both gives benefits and suffering. It has a low wingloading which enables it to fly at lower A.o.A. at lower speeds, - A.o.A. creates drag.
Hence the maneuverability in the low speed region as well as good climb.
Yep, as I understand it. In the low speed realm even a small advantage becomes HUGE. At higher speeds the force of induced drag is hardly a factor.
http://www.av8n.com/how/htm/4forces.html#fig-power-ias
As you go slower and slower, induced drag increases dramatically and parasite drag becomes almost negligible.
Notice the shap of the curves for the power of induced drag. It stays pretty much steady until airspeed drops to a certain point. Get slow enough and the wall comes up!
Crumpp
-
Crumpp,
Don't create theories without understanding, you can't say anything about lift distribution based on pressure distribution of one wing side only.
All we can see is that the Fw 190 has a relatively bit thicker and wider profile in the wing tip. This is natural outcome of the wing shapes.
gripen
-
Don't create theories without understanding, you can't say anything about lift distribution based on pressure distribution of one wing side only.
It's in black and white Gripen. Just read the link. Besides we saw your understanding by creating theories off Lednicer's lift distribution as you came up with gooble-dee gook to solve for the wingtip efficiency off of it!!
I had no wind tunnel data to calculate e factor of the Fw 190. But Lednicer calculated span loadings and here is estimated e factors calculated by taking 8 samples (semispan fractions 0,2-0,9) and measuring distance to the elliptical span loading:
If this guy can't figure it out from Span Loadings:
http://www.grc.nasa.gov/WWW/K-12/airplane/benson.html
I highly doubt you can either. So don't hop on your high horse about creating theories. You throw your fair share around.
Crumpp
-
Crumpp,
Well, all I see here is that you see something like an elliptic shape in the pressure distribution picture and somehow you conclude that it tell something about lift distribution.
Regarding estimating e factor from the lift distribution. We know that there is a strong correlation between relative lift distribution and e factor. And that is what my calculations show. Only suprising thing is the accuracy, apparently rise of the viscous drag correlates strongly with the rise of the induced drag
gripen
-
Well, all I see here is that you see something like an elliptic shape in the pressure distribution picture and somehow you conclude that it tell something about lift distribution.
It's not really what you see Gripen. It's more like what you do NOT see in the Spitifire. Obviously the "theoritically perfect" ellipse you give the Spitfire credit for having does not exist.
And that is what my calculations show.
Yes, Heavily slanted towards you favourite plane and performed with pseudo-logic.
Crumpp
-
Originally posted by Crumpp
It's not really what you see Gripen. It's more like what you do NOT see in the Spitifire. Obviously the "theoritically perfect" ellipse you give the Spitfire credit for having does not exist.
I see the the same thing as Lednicer and everybody else:
(http://personal.inet.fi/koti/soon.moro/sp.jpg)
The lift distribution of the Spitfire "is not elliptical, though it is probably the most optimum of the three from the induced drag standpoint".
And the wind tunnel data and Fw data confirm this:
P-51 e=0,77
Fw 190 e=0,78
Spitfire e=0,88
And as noted several times you are the only one to claim (http://www.hitechcreations.com/forums/showthread.php?s=&postid=1368542#post1368542) e factor value 1 for the Spitfire.
Originally posted by Crumpp
Yes, Heavily slanted towards you favourite plane and performed with pseudo-logic.
Well, if someone is heavily slanted towards something in this thread, it's you.
And your logic is generally laughable; your theories on swept wing and pressure distribution as seen above are good examples.
gripen
-
P-51 e=0,77
Fw 190 e=0,78
Spitfire e=0,88
Oh Yeah! These are the values calculated from HOW again?
I had no wind tunnel data to calculate e factor of the Fw 190. But Lednicer calculated span loadings and here is estimated e factors calculated by taking 8 samples (semispan fractions 0,2-0,9) and measuring distance to the elliptical span loading:
Off the "span loading"....
That is quite the feat of calculator jockying. You should publish a paper on your technique and inform the rest of the aeronautical engineering world. No trained aeronautical engineer seems to be able to do it.
They all say the same thing. You can tell the wings had some washout but it does not tell you much about wingtip efficiency.
Crumpp
-
Originally posted by Crumpp
Oh Yeah! These are the values calculated from HOW again?
As shown above and in the Perkins&Hage.
Originally posted by Crumpp
That is quite the feat of calculator jockying. You should publish a paper on your technique and inform the rest of the aeronautical engineering world.
The rest of the world all ready knows that the lift distribution and the efficiency factor have a strong correlation.
Besides these values are calculated from wind tunnel data and Fw data not from lift distribution chart, still the estimates from the lift distribution chart are suprisingly accurate as seen above.
gripen
-
As shown above and in the Perkins&Hage.
Show us in Perkins & Hage were you can calculate wingtip efficiency factor from span loading.
What page?
Besides these values are calculated from wind tunnel data and Fw data not from lift distribution chart, still the estimates from the lift distribution chart are suprisingly accurate as seen above.
Surprisingly accurate??
:eek:
Crumpp
-
Originally posted by Crumpp
Show us in Perkins & Hage were you can calculate wingtip efficiency factor from span loading.
I say above that these values are calculated from the wind tunnel data and the Fw data, not from span loading. The system is described in the Perkins&Hage (p.90-95)
Originally posted by Crumpp
Surprisingly accurate??
The values of e calculated from the wind tunnel data and Fw data:
P-51 e=0,77
Fw 190 e=0,78
Spitfire e=0,88
The values of e estimated from the span loading chart:
P-51 e=0,80
Fw 190 e=0,79
Spitfire e=0,86
gripen
-
Gripen,
What is the setup for the Spitfire in your drag polar? Is it a wooden model, actual plane, and if so what varient?
The values of e calculated from the wind tunnel data and Fw data:
Don't match the aeronautical engineers calculations of .87 e factor for the same data for the FW. I would say something is off in yours.
Crumpp
-
I have touched the dihedral (somewhere in some thread) as a possible factor.
Why?
It's bound to reduce lift. A dihedral of 90 degs would for instance produce no upwards lift at all.
It's basically a stability enhancer, however it must effect lift, and therefore drag, be it little or not.
At the wingtips one has upwash, right outside of it. This will probably give the tip a bit of a push upwards.
(what turned out to be strong enough to bend the Hurricane's wings upwards when dived to the limit)
So I'm wondering where this would come into the equation.
For some reason, a typical dihedral is somewhere between 4% and 6%.
Any comments?
-
Originally posted by Angus
I have touched the dihedral (somewhere in some thread) as a possible factor.
Why?
It's bound to reduce lift. A dihedral of 90 degs would for instance produce no upwards lift at all.
It's basically a stability enhancer, however it must effect lift, and therefore drag, be it little or not.
At the wingtips one has upwash, right outside of it. This will probably give the tip a bit of a push upwards.
(what turned out to be strong enough to bend the Hurricane's wings upwards when dived to the limit)
So I'm wondering where this would come into the equation.
For some reason, a typical dihedral is somewhere between 4% and 6%.
Any comments?
As I recall from ground school, dihedral enhances directional stability. If a sideways air load strikes the top of one wing, it must strike the bottom of the other. This rolls the plane into the sideways load, maintaining direction. Another commonly used explanation is that when the plane rolls, the lift of the high wing decreases as the lift of the low wing increases--at least until the low wing becomes horizontal, at which time the high wing lift decreases faster than the low wing's lift. The high wing will always have a steeper angle to the horizon, you see.
The wingtip vortices are ALL drag--any lift they (arguably) provide is overshadowed by the huge drag penalty--remember that moving air is work (in the sense of basic physics) and work requires energy. Many modern planes have wingtip end plates to decrease the amount of spanwise spillage--the actual cause of vortices. Theoretically,this reduces drag and also decreases the strength of the vortex that has been blamed for so many "clear air turbulence" accidents when light planes follow heavy planes too closely.
According to mr. trigonometry, by the way, dihedral should have a very small effect on lift. Doing the math, a 4 degree dihedral would decrease lift by 7% in straight, level flight.
shubie
-
Nice! :)
So if there is a force acting to push the tip upwards, it is wasted energy right?
-
Originally posted by Angus
Nice! :)
So if there is a force acting to push the tip upwards, it is wasted energy right?
No, but if there is AIRFLOW upward at the tip--caused by high pressure air under the wing trying to go to the low pressure above the wing--that is work being done by the wing that doesn't help the plane fly--the very defintion of drag. Pushing the tip up is lift, if it isn't BENDING the tip or the wing. If the tip bends, it is (again) drag.
shubie
-
Speaking of washout, could somebody explain to me the theory behind it ? Or point towards a link about it ?
-
It's basically the air washing outwards towards the end of the wing.
I'm sure that rshubert can word it better, but that's basically it.
Regards
Angus
-
Originally posted by Angus
It's basically the air washing outwards towards the end of the wing.
I'm sure that rshubert can word it better, but that's basically it.
Regards
Angus
Washout is a deliberate bending of the wingtip downward (chordwise) so that the angle of attack at the wingtip is less than the angle of attack at the wingroot. The idea is to get the wingtip to stall AFTER the wingroot, so that you have aileron control right up to the stall. If the root stalls first, the plane will have less tendency to drop a wing at the stall, since the pilot will be able to apply aileron to counteract the roll induced by torque, prop wash, and/or P factor. A similar effect can be seen via the use of fixed slots, like we see on the Me-163 Komets. They "blow" air over the upper wing surface, maintaining smooth airflow over the wing at a high(er) angle of attack. I don't know if the Komet has any washout in combination with the slots.
I think washout is only a couple of degrees, usually. And it is factored into the lift calculations for a particular airfoil design.
shubie
-
So, if you have a dihedral, you have to increase washout right?
rshubert, either you are very well read, or a pro, must be.
I am honoured to be exchanging words with you M8
-
Originally posted by Crumpp
What is the setup for the Spitfire in your drag polar? Is it a wooden model, actual plane, and if so what varient?
It's a model. I have given the source above, get it if you are interested.
Originally posted by Crumpp
Don't match the aeronautical engineers calculations of .87 e factor for the same data for the FW. I would say something is off in yours.
The Fw data gives value 1,24 for K at climb flying condition, that results e factor value 0,78. There is no other result with this value of the K.
gripen
-
It's a model. I have given the source above, get it if you are interested.
Yes I am interested. You do know that a model is not worth comparing to the real plane. It would be interesting to compare with the FW-190V5g model test's. Pointless, but interesting.
Crumpp
-
How do you define the E factor in maximally effective english anyway?
Not kidding :p
-
There is ample evidence that the Spitfire wing was reletively hard to make. The spar itself required very precise machining of several interlocking tubes. Supermarine had a good deal of trouble sub-contracting the work.
One would not choose the Spitfire layout as an economical way to enter into mass production of fighter aircraft in 1940.
-blogs
Originally posted by gripen
Seems that Mitchell ended to the elliptic form for several reasons and need to place the wheels, guns and ammunition inside wing are probably among these. Otherwise, it's quite probably that they did not think mass production at all; in the mid thirties very few types reached large production numbers.
Regarding stall, the Spitfire had generally good stall characters like the Bf 109 ie no tendency to tip stall.
gripen
-
Originally posted by Crumpp
You do know that a model is not worth comparing to the real plane.
I don't know any particular reason why the wind tunnel results should not be comparable; basicly it's size of the model which changes and how clean the model is. The RAE data is for the Spitfire I and Allison Mustang, both are 1/6 scale models. I don't know how the Fw data is generated but the dataset contains drag data for separate parts of the airframe so it might be from smaller size model test too.
In most cases the wind tunnel data seems to be quite comparable with other wind tunnel tests but the flight tests seem to give lower values.
gripen
-
In most cases the wind tunnel data seems to be quite comparable with other wind tunnel tests but the flight tests seem to give lower values.
Windtunnel data on a model has little to do with the actual aircraft with the exception of making general assesments. You certainly cannot take exact values off it. Ask Hitech.
You can say:
This design will fly and be stable.
This design will be "X" fast in level flight
etc..
No whether they hold true or on can only be determined by the actual aircraft. Drag data would have to come from an actual aircraft in order to have any kind of meaning.
The RAE data is for the Spitfire I and Allison Mustang, both are 1/6 scale models.
Completely worthless for comparision. We could compare the wooden model of the FW-190 with them but that would be a complete waste of time as well.
In most cases the wind tunnel data seems to be quite comparable with other wind tunnel tests but the flight tests seem to give lower values.
Ahh, that explains a lot.
Crumpp
-
Since when did wind tunnels get such a bad rep?
Sure you have to be careful with design, reynolds numbers, etc, but most of these things had been worked out for sub-sonic flows well before the war.
The advantage of a wind tunnel test over flight tests is that the experiment can be better controlled and experiments can be more easily compared with other experiments. The disadvantage is that wind tunnel models may not capture all the features that will affect an actual production model...
-blogs
Originally posted by Crumpp
Windtunnel data on a model has little to do with the actual aircraft with the exception of making general assesments. You certainly cannot take exact values off it. Ask Hitech.
You can say:
This design will fly and be stable.
This design will be "X" fast in level flight
etc..
No whether they hold true or on can only be determined by the actual aircraft. Drag data would have to come from an actual aircraft in order to have any kind of meaning.
Completely worthless for comparision. We could compare the wooden model of the FW-190 with them but that would be a complete waste of time as well.
Ahh, that explains a lot.
Crumpp
-
"In most cases the wind tunnel data seems to be quite comparable with other wind tunnel tests but the flight tests seem to give lower values. "
I guess that is because there are many minor factors which may have a significant effect combined. Surface design, exhaust system vs. exhaust thrust, cowling fitting, temporal jig misalignment causing badly formed fuselages etc.
I'd imagine that eg. the air that enters the engine cowling of FW is problematic from aerodynamic modelling point of view. Do they consider the air intake to be closed or to flow through to some extent or what. Considering the size of the air intake it may have a significant effect on how well the model correlates with the real life a/c.
:confused:
-C+
-
Hm... Generally the effect of the powerplant (slipstream, radiators, heat effects etc.) is a unknown factor in the all wind tunnel data and it's a problem when testing a scale model or a full size aircraft in the wind tunnel. But otherwise I see wind tunnel as a good and practical way to study planes.
gripen
-
I agree. I'd imagine that the wind tunnel gives a rather accurate estimate of the behaviour of the aerodynamics and good results when comparing similarly manufactured scale models. The error estimate calculations may be tricky, though.
Otherwise in my opinion a combat aircraft and their abilites cannot be very accurately evaluated by the wind tunnel data alone even if it gives a good estimate of the stall speeds and stall and drag behaviour of the airframe.
I thought the Lednicher article to be a bit torso on part of 190. I dont think he is aware of the possibility of bad aileron tuning which could cause similar effects as torsional warping of the wing in tight turns. Based on the NACA report on roll rate I think is rather unlikely that 190 had a torsionnally loose wing considering its excellent roll rate which would not be possible if his assumption is true. :confused: Or is it? Dunno for sure.
-C+
-
But otherwise I see wind tunnel as a good and practical way to study planes.
It definately is and I am not saying wind tunnel data is useless for aeronautical engineers or for drawing general conclusions off a design.
For comparision purposes windtunnel data off of a model and not the actual airplane is worthless for comparision with the real plane. Not the general conclusions but when you try and get exact figures it most certainly is and can lead to wrong comparison's between different aircraft designs.
In my research I made a similar mistake with FW-190V5g. V5g is the big wing test prototype with a BMW-801 engine. Very similar to the production model of the FW-190A. The engineers who are helping me out immediately pointed out that the actual data is worthless for direct comparision's with other aircraft. Hitech confirmed this when I offered the report to them as well.
That is why I hunted down some windtunnel test's of the actual aircraft, made the trip, and spent the money to get them. Those drag polars would be useful for comparing drag polars from another actual plane but not a model.
Comparing models is just that. Comparing a model and not the real thing. The FW-190 report "Drag data for Aircraft" comes from actual aircraft as well.
I guess that is because there are many minor factors which may have a significant effect combined. Surface design, exhaust system vs. exhaust thrust, cowling fitting, temporal jig misalignment causing badly formed fuselages etc.
That is absolutely correct as I understand it. Even with the actual aircraft you have to be careful. Just look at the P51 wind tunnel test reports. Several of them were done with a preproduction model that did not have simple things like exhaust stacks mounted. These simple things add up and this is why you see some claims of extremely low drag on the 51. It was a low drag design but not nearly as low as some of the claims.
The intakes are problematic for any model. In the case of the FW-190A, since plane itself is such a small physical size, the intake area is very compariable to the Spitfires radiators. Just a different shape.
Crumpp
-
I thought the Lednicher article to be a bit torso on part of 190. I dont think he is aware of the possibility of bad aileron tuning which could cause similar effects as torsional warping of the wing in tight turns. Based on the NACA report on roll rate I think is rather unlikely that 190 had a torsionnally loose wing considering its excellent roll rate which would not be possible if his assumption is true. Or is it? Dunno for sure.
One thing for sure Charge. The FW-190 had extremely sensitive ailerons that needed precise adjusting. Additionally according to both the pilots who flew it and the manuals, the FW-190 had multiple hinges to "tune" the ailerons to a desired performance band. The hinges adjusted the height of the aileron in it's mounting giving the pilot different "max roll" performance at different parts of the envelope.
If I had to say the achilles heel of the 190, it would be it's aileron adjustments. They are simple machined blocks of duraluminum with grooved slots for the mounting bolts. These blocks fit on top of the aileron mounting block in the wing which is simply a fixed flat block. The aileron is mounted (with whatever hinges) on this flat block and set in the correct position. The bolts are then tightenend down and they are the only thing holding the aileron in it's adjustment point. I imagine those blocks could slide and the aircraft required constant inspection of the adjustment.
The aileron adjustment had severe effects on the handling qualities of the design. I know that the FW-190A5 the USN tested had out of adjustment ailerons. When you read the Luftwaffe caution statements, advisory reports, and Focke-Wulf data it becomes obvious.
Being that the main wing spar is solid on the FW-190 and runs the entire length of the wing underneath the fuselage, I imagine it torsionally strong. I have the complete report of RAE 1231 which the NACA graph is based on. In it the test pilot states the FW-190 flown has the stiffest ailerons of the three FW-190's he has flown. Additionally with Frise alerions you are going to get a rather wide performance band, AFAIK.
Crumpp
-
"The intakes are problematic for any model. In the case of the FW-190A, since it is such a small physical size, the intake area is very compariable to the Spitfires radiators. Just a different shape."
Edit: Different shape yes, but also probably causing rather much drag in the process, too, by increasing the face area of the a/c.
Yes, but their placement may have effect on other factors again, as noted of the Spitfires radiator which, AFAIK, had somekind of effect on wing lift. Don't remember exactly if it was considered good or bad or insignificant. And to me it seems that it has quite a bit effect if the radiator is able to close completely or not.
I have never understood where the cooling air exits in 190A. In first models the only vents in the cowling were for the exhaust tubes and because of overheating the cooling louveurs were made in the aft cowling to allow exit of hot air. To me it seemed like a bit idiotic design flaw to leave them out in the first place...
Maybe that centralized drag was one of the factors which gave the 190 such a good rollrate as there was no off-balancing drag effect anywhere else but just around the thrust line vector (ie. engine camshaft?) And maybe that is the reason they decided to use similar cowling design in D-models. So I'm guessing that if they had made a different eg. P51 style radiator on D9 the drag could have been less, but the roll rate could have suffered too?
-C+
-
Originally posted by Crumpp
Comparing a model and not the real thing.
Well, as noted above, the Fw data seems to be based on wind tunnel data possibly with some kind of model.
Besides I don't see any reason why the data determined with a model in the wind tunnel should not be comparable with the data determined with a full size aircraft in the wind tunnel.
Regarding ailerons, the Brits captured at least three Fw 190s intact (landed intact to the British air fields for various reasons) so the adjustment of the ailerons is very unlikely the reason for tendency to stall from the turn. But that is not a subject of this thread.
gripen
-
Yes Gripen, I believe the Lednichers article on that part that the wing profile was prone to a sudden stall but when it is considered a design flaw despite washout and all I'm bound to think that the ailerons where not properly adjusted on that particular aircraft Brown flew.
Sorry 'bout this off current in your thread on my part Gripen.
-C+
-
Originally posted by Angus
So, if you have a dihedral, you have to increase washout right?
rshubert, either you are very well read, or a pro, must be.
I am honoured to be exchanging words with you M8
Not a professional--just an electrical engineer who flies for fun. As to the question, I don't know. It doesn't necessarily follow, does it? The stall is purely an angle of attack question, and the washout affects the wingtip angle of attack. Dihedral doesn't affect angle of attack, so I guess that playing around with dihedral would not have any effect on the amount of washout needed to provide control on the edge of a stall. But that is a GUESS, not a known fact.
It is an interesting question, though--and points out that the subject of wing design is not easily explained in terms of one or two factors. NACA (NASA's predecessor), the Brits, the Germans, and everybody else spent a LOT of time on wing designs during the '20s and '30s, many of which are still in use today.
shubie
-
Originally posted by rshubert
Dihedral doesn't affect angle of attack, so I guess that playing around with dihedral would not have any effect on the amount of washout needed to provide control on the edge of a stall.
Yep, extreme amount of dihedral "might" have a similar effect as swept back wing due to increased downwash at low speeds but generally it's there just for stability.
gripen
-
Edit: Different shape yes, but also probably causing rather much drag in the process, too, by increasing the face area of the a/c.
I don't think they would cause anymore drag than a radiator. In fact I would think less. Drag is caused by surface area, correct? Sure cooling fins on an air cooled radial are there to increase the surface and would increase drag. Compared to radiators which are fairly large as were the Spitfire Mk IX's, I don't see that the difference is that great.
In fact Lednicer seems to be correct. If the FW-190 was as big as the spit it would have more drag. It's smaller physical size though gives it less form drag than the Spitfire.
I have never understood where the cooling air exits in 190A. In first models the only vents in the cowling were for the exhaust tubes and because of overheating the cooling louveurs were made in the aft cowling to allow exit of hot air. To me it seemed like a bit idiotic design flaw to leave them out in the first place...
Air can exit around the exhaust stubs. Got to agree though. The space in not very big and someone was surely being overly optimistic that the small volume exiting the space would carry off the excess heat.
Got a nifty report examining different ways to seal the cowling.
They were going for minimum drag on the cowling job and tried to get away with as little excess space under it as possible. What I don't understand is why they put them on the inside at first in the FW-190A1-A4. The FW-190A5 and later gills are louvers on the outside and are much more aerodynamic. They work better too. Putting the louvers on the inside makes no sense to me! Your trying to get the air out and the slit for the gills is certainly not aerodynamic.
Regarding ailerons, the Brits captured at least three Fw 190s intact (landed intact to the British air fields for various reasons) so the adjustment of the ailerons is very unlikely the reason for tendency to stall from the turn.
The Geschwaders had a hard time keeping the ailerons adjusted. Why would the RAF be any better at it?
Besides I don't see any reason why the data determined with a model in the wind tunnel should not be comparable with the data determined with a full size aircraft in the wind tunnel.
Give Pyro a call and ask him this.
Crumpp
-
Originally posted by Guppy35 .Keep in mind the 190s were actually Rumanian IAR80s
The FW190 is not a Romanian IAR 80.
-
Originally posted by Crumpp
In fact Lednicer seems to be correct. If the FW-190 was as big as the spit it would have more drag.
Not a subeject of this thread but Lednicer's numbers give 3,5% difference in drag and the values for the Fw 190 are based on manufacturers data. Basicly a neglible difference.
Originally posted by Crumpp
The Geschwaders had a hard time keeping the ailerons adjusted. Why would the RAF be any better at it?
Not a subject of this thread but my understanding is that this theory is borne out in the same place as "wet lifting area", strange theorie on swept back wings and pressure distribution etc...
Start a new thread and show the physical connection between stall in the turn and the ailerons.
Originally posted by Crumpp
Give Pyro a call and ask him this.
Pyro is most wellcome to join in discussion any time.
Basicly you are changing your argument all the time depending on which seem to support your agenda on the Fw 190. In the beginning of this thread you were arguing that some aspect ratio based formulas give correct value of e. Then you created various strange theories on lift distribution. Now you claim Fw data flight tested despite it's very unlikely. Still, you have not been able to bring in any evidence nor calculate something properly so far.
gripen
-
Basicly you are changing your argument all the time depending on which seem to support your agenda on the Fw 190.
BS Gripen.
I thought you knew about aeronautics to know better. I assumed you were smart enought to use good data in the first place. This entire argument would have never happenend had I known your drag polars came off of 1/6 wooden models and not an actual aircraft.
Not a subeject of this thread but Lednicer's numbers give 3,5% difference in drag and the values for the Fw 190 are based on manufacturers data. Basicly a neglible difference.
Was not neglible when you were claiming it was not true.
Not a subject of this thread but my understanding is that this theory is borne out in the same place as "wet lifting area", strange theorie on swept back wings and pressure distribution etc...
Yeah, OK. Guess the Luftwaffe just killed trees by writing memo's, manuals, and tons of reports detailing the importance of properly adjusted ailerons to turning performance and stall.
Still, you have not been able to bring in any evidence nor calculate something properly so far.
Lets see and compare total drag @ 440fps at the Spitfires FTH altitude with the correct data:
BF 274 - 795.641828 @ 440 fps
Total Drag FW-190A5 -718.2845242@ 440 fps
BS354 - No data listed except the A/C was 7 mph slower w/50 bhp more output than the Merlin 66 @ (+18)
BS543 - 810.863789
FW-190A5 - 705.7484202
BS551 - 758.4823884
FW-190A5 - 704.890256
EN524 with 4 bladed prop - 765.4635575
FW-190A5 - 702.4712376
BS310 with 4-bladed prop - 793.0539968
FW-190A5 - 751.8834689
BS310 with 5 bladed prop - 794.0663191
FW-190A5 - 749.7161496
JL165 - done to death
MA 648 Merlin 66 (+18) SU Pump - 838.0549776
FW-190A5 - 804.8037275
BS310 - 792.6552588 Parasitic drag - 691.672177
FW-190A8 - 794.3964059 Parasitic drag - 606.4158622
At 315mph the total drag situation changes completely in the FW-190A8 favor.
JL165 vs FW-190A8 has been done to death, again in the FW-190A8's favor.
MA 648 is the ONLY Spitfire Mk IX to beat the FW-190A8 for drag and here is why:
quote:
--------------------------------------------------------------------------------
External equipment of the above four aircraft was similar with the exception that MA.648 had the new pattern of air intake.
--------------------------------------------------------------------------------
MA648 - 837.1701001 parasitic drag - 751.6212076
FW-190A8 - 869.2121249 parasitic drag - 709.9624067
The parasitic drag is always in the FW-190's favor but the total drag drops to within a few pounds of each other at 315 Mph but never swings in the FW-190's favor.
So ONE example of the Spitfire Mk IX with an experimental air intake beats the FW-190A8 in total drag. The others are behind their FW-190A contemporary the whole way.
That about covers the Spitfire Mk IX's vs. their contemporary FW-190 adversary.
Crumpp
-
Originally posted by Crumpp
This entire argument would have never happenend had I known your drag polars came off of 1/6 wooden models and not an actual aircraft.
Well, you have a very short memory, until last days you have been arguing that generalized formulas give accurate results.
Please tell me what's wrong with scale models and why the results should not be comparable the with the full scale model in the wind tunnel.
Besides, as noted above the Fw data seems to be based on scale models as well. There was no large wind tunnel in the Germany and the document is dated december 1944 as well as plane variants listed. In addition the datasheet gives detailed drag data and
claims that drag data includes interference drag, these are typical things for model testing.
Originally posted by Crumpp
Yeah, OK. Guess the Luftwaffe just killed trees...
Start a new thread and post your evidence there and also learn to calculate first; you still can't calculate drag...
And most important: Read the forum rules.
Originally posted by Pyro
2- Threads should remain on topic, do not "hijack" topics.
gripen
-
In addition it should be noted that the Fw datasheet claims:
"Oberflächengüte einschl."
Which means that surface quality is supposed to be as perfect as possible. Another clear indication of wind tunnel data.
gripen
-
Hello guys.
The drag of those two seems to fall short of each other, the 190 being smaller and cleaner if anything, and note, that in the real world it had a better finish.
What Crumpp said:
"The parasitic drag is always in the FW-190's favor but the total drag drops to within a few pounds of each other at 315 Mph but never swings in the FW-190's favor. "
I think this pretty well spells it out. The 190 has less parasite drag, but more induced drag due to higher A.o.A. in the flight.
Once the Spitfire was "cleaned up" a bit and boosted upwards, it outperformed the 190A.
Such as the Spit VIII.
Am I right?
(Noooo, you're not both gonna say I'm wrong are you :D )
-
"Once the Spitfire was "cleaned up" a bit "
What do you mean? Adding another huge radiator under the wing?
The Griffon cowling was more streamlined, though, and the IX had more pointed spinner but still rather high engine profile if we compare those...
-C+
-
Nonono, not talking about Griffon Spits.
There were modifications withing the Mk IX series, and then there was the superb Mk VIII.....
-
Which means that surface quality is supposed to be as perfect as possible. Another clear indication of wind tunnel data.
As opposed to a laquered wooden 1/6th scale Model!!
OMG Gripen. You have dropped off the deep end in your defense.
Crumpp
-
Please tell me what's wrong with scale models
http://www.hitechcreations.com/htcindex.html
Give em a call, I am sure Pyro will help you out.
Crumpp
-
Originally posted by Crumpp
As opposed to a laquered wooden 1/6th scale Model!
The Fw data is probably determined with even smaller model (Ta 152s).
Originally posted by Crumpp
Give em a call, I am sure Pyro will help you out.
Pyro is most wellcome to join in discussion anytime.
gripen
-
Originally posted by Angus
Hello guys.
The drag of those two seems to fall short of each other, the 190 being smaller and cleaner if anything, and note, that in the real world it had a better finish.
I have added flat plate areas for Cd0 to the another thread.
gripen
-
There was no large wind tunnel in the Germany
Totally silly statement. You need to research the Herman-Goering test facility.
Crumpp
-
Originally posted by Angus
Nonono, not talking about Griffon Spits.
There were modifications withing the Mk IX series, and then there was the superb Mk VIII.....
Huh? Charge is right...aerodynamic cleanup on Spitfires during the war? When, what? IMHO that was the greatest failing of the whole series, ABSOLUTELY nothing was done, they started with a nice clean aircraft, then they added cannon bumps up and down, horrific Griffon engine bumps, landing gear bumps, adding another huge and ugly radiator under the wing with the MkIX/VIII, and making them even bigger on the XIV.. I`ll never understand why such a complete lack of improvement in the airframe`s aerodynamics... they could EASILY made the plane 30-40mph faster...
Oh, pardon, I am wrong. They managed to make the the tail wheel retractable on ca. 10% of all Spits produced. Most didn`t even had that... That added about 4mph... :D
-
Originally posted by Crumpp
Totally silly statement. You need to research the Herman-Goering test facility.
Ah, another new historical fact. Please tell us more.
Generally
-
Originally posted by gripen
Ah, another new historical fact. Please tell us more.
Hmm, indeed the lack of large wind tunnel facilities in Germany DOES sound *strange*. From memory, they used among others the Peenemunde wind tunnels for planes, which were origibuilt for testing the V-2s at high Machs.
-
AFAIK there were three large wind tunnels around in forties: Tsagi, NACA and Chalais-Meudon.
I'm most interersted to hear about others. There were advanced wind tunnels in the germany but I'm not aware if any of these were big enough to take fighter size aircraft. As and example Chalais-Meudon was 16x8m.
gripen
-
I'm most interersted to hear about others. There were advanced wind tunnels in the germany but I'm not aware if any of these were big enough to take fighter size aircraft. As and example Chalais-Meudon was 16x8m.
I have some good pics of a 109 hanging in the Herman Goering. I will post it when I get home. It is definately NOT a model.
Crumpp
-
I looked up it at NACAs archieves, fairly good stuff to read. Didnt read the whole section, but basically it says that most modern wind tunnels mimic the one built during WW1 at Gottingen, and the large ones follow the example of the 2nd generation large wind tunnel at Gottingen, Germany in 1916, again by Herr Prandtl. But I did not find any other *very large* windtunnels big enough to take wholesized aircrafts, even the NACAs own, while being so big that it was capable of taking bombers, was limited to slow flight speed emulations, and so they used it for t/o and landing tests.. I may have missed it, but the others were only capable of aircraft parts tests, not complete a/c. Some details I found on (some of?) the Peenemunde windtunnels, one(?) of the m being described as a rather small (half meter big) w/t, but which was capable of simulating Mach 5!
Good read... http://www.hq.nasa.gov/office/pao/History/SP-440/contents.htm
-
(http://images.amazon.com/images/P/0300063679.01._SCLZZZZZZZ_.jpg)
The Peenemünde Wind Tunnels
A Personal Memoir
Peter P. Wegener
In 1943, on orders from the German Air Ministry, young physicist Peter P. Wegener left the Russian front and reported to the Baltic village of Peenemünde. His assignment was to work at the supersonic wind tunnels of the rocket laboratories of the German Army. Here Wernher von Braun led a team that developed the V-2, the world’s first large rocket-powered guided missile, and laid much of the groundwork for postwar rocket development.
In this fascinating book, Wegener recounts his experiences during Hitler’s time, World War II, and his years at Peenemünde. He tells how he was working one night in August 1943 when the Allies bombed the laboratories, but left the wind tunnels undamaged. The tunnels were moved to Bavaria, and Wegener was ordered to follow in 1944. After the war, the tunnels were moved again--this time to the United States, accompanied by the author and other German scientists. Shortly before the end of the war, Wegener visited Germany’s underground V-2 production plant to retrieve archival material on aerodynamics that had been stored in caves for safekeeping. He describes the appalling history of the concentration camps where SS guards watched over inmates who toiled underground in inhuman conditions and often did not survive. A photoessay enhances this remarkable memoir.
There is a photo of a 109E-3 in the wind tunnel of the Goring Aviation Resaerch Institute in Volkenrode near Brunswick, pg 126 of Bf109 A-E by Radinger/Schick.
-
There is a photo of a 109E-3 in the wind tunnel of the Goring Aviation Resaerch Institute in Volkenrode near Brunswick, pg 126 of Bf109 A-E by Radinger/Schick.
Yep, that would be this one:
(http://www.onpoi.net/ah/pics/users/503_1100816162_109windtunnel.jpg)
The Fw data is probably determined with even smaller model (Ta 152s).
Is a totally false statement. The FW Data comes off the aircraft. Not a model.
Crumpp
-
AFAIK that is a 8m tunnel called A3. Not large enough for lift coefficient measurements.
gripen
-
"There were modifications withing the Mk IX series, and then there was the superb Mk VIII....."
The VII was stretched a bit and was mounted with 61, 64 or 71 Merlin and was the first to have those two similar radiators under the wings and it had a retractable tail wheel. The VIII was unpressurized variant with Vokes filter with a few different engines.
The tail wheel was however made fixed in IX and the late IX model had again the retractable one where as the other later models again had fixed wheel and again from XII onwards the tailwheel was retractable. Maybe just a feature of parallel development, dunno.
-C+
-
Many things getting developed at the same time, definately.
Never heard of a Mk VIII with a Vokes filter, can you give some more data on that please? :)
-
As if the 109 did not have protrusions, Issy.:rolleyes:
Bulges for the cowl mgs, bulges for the wheels, numerous other little bulges and scoops, an air intake sticking out from the side of the fuselage, 3 radiators (2 coolant, 1 large oil), ADF loop, a retractable tail wheel that was sometimes locked down, extended tail wheel strut, gondola guns hung from the wings, . . . . . . .
-
Originally posted by Crumpp
The FW Data comes off the aircraft. Not a model.
The 8m A3 was the largest wind tunnel in the LFA and the wingspan both Ta 152s is more than 10m, even the the short span Bf 109 V24 was too large for the A3, that's why it was tested later in the Chalais-Meudon.
Besides, most of the high speed testing in the LFA was done in the 2,8m tunnel called A2.
And you should still point out what is wrong with the testing with models.
gripen
-
And you should still point out what is wrong with the testing with models.
Did not really press the engineers about it Gripen. They told me you can't do it and Pyro told me you can't do it. I did not argue and gave them data off of actual aircraft.
The 8m A3 was the largest wind tunnel in the LFA and the wingspan both Ta 152s is more than 10m, even the the short span Bf 109 V24 was too large for the A3, that's why it was tested later in the Chalais-Meudon.
Wingspan on that E3 is 9.9 meters. Looks like it has some room to spare Gripen. At least in the photo.
For the FW-190A8 Finish. This came out of the original report translated in Valkyrie Publications FW-190A8 Pilots Manual. It is one of several errors in that book.
(http://www.onpoi.net/ah/pics/users/503_1100818835_fw-190a8graph2.jpg)
It clearly states the finish of the aircraft is "Surface primed and Smooth Painted" Smooth painted refers to standard RLM semi-gloss aircraft paints.
Crumpp
-
Ok so the topic is about parasitic drag. How much drag is invoved here?:lol
(http://www.1077thebone.com/jocks/lnt/goofy/images/bigstealthchick.gif)
-
Nice arms on that chick...
:(
Crumpp
-
What's that thing under that thing?
-
The f4f was cleaned up on the basis of full scale wind tunnel tests. NACA made quite a business of this starting in the late 1930s.
-Blogs
Originally posted by Kurfürst
I looked up it at NACAs archieves, fairly good stuff to read. Didnt read the whole section, but basically it says that most modern wind tunnels mimic the one built during WW1 at Gottingen, and the large ones follow the example of the 2nd generation large wind tunnel at Gottingen, Germany in 1916, again by Herr Prandtl. But I did not find any other *very large* windtunnels big enough to take wholesized aircrafts, even the NACAs own, while being so big that it was capable of taking bombers, was limited to slow flight speed emulations, and so they used it for t/o and landing tests.. I may have missed it, but the others were only capable of aircraft parts tests, not complete a/c. Some details I found on (some of?) the Peenemunde windtunnels, one(?) of the m being described as a rather small (half meter big) w/t, but which was capable of simulating Mach 5!
Good read... http://www.hq.nasa.gov/office/pao/History/SP-440/contents.htm
-
Originally posted by Crumpp
Did not really press the engineers about it Gripen. They told me you can't do it and Pyro told me you can't do it. I did not argue and gave them data off of actual aircraft.
Well, the problem is you have claimed things which are not true several times in this thread. I don't see here any evidence that Pyro or somebody else has really said so. And if somebody has said so, I'd like to know why?
Probably over 90% of all wind tunnel testing has allways been done with models and I really like to hear what's wrong with that.
Originally posted by Crumpp
Wingspan on that E3 is 9.9 meters. Looks like it has some room to spare Gripen. At least in the photo.
You can easily see from the photo that the wing tips are out of the tunnel, it's the 8m tunnel A3. AFAIK the Bf 109 V24 had short span wing, less than 9m, at least that is the configuration it was tested in the Chalais-Meudon.
Originally posted by Crumpp
For the FW-190A8 Finish...
Not the topic of this thread and learn German...
gripen
-
Angus, the VIII which where use outside Europe featured a well streamlined Vokes filter which was a much better design than the earlier Vokes or Aboukir filters seen in SpitV. (Aboukir was in fact claimed to be quite good with only 5mph speed loss...)
-C+
-
Originally posted by MiloMorai
As if the 109 did not have protrusions, Issy.:rolleyes:
Bulges for the cowl mgs, bulges for the wheels, numerous other little bulges and scoops, an air intake sticking out from the side of the fuselage, 3 radiators (2 coolant, 1 large oil), ADF loop, a retractable tail wheel that was sometimes locked down, extended tail wheel strut, gondola guns hung from the wings, . . . . . . .
The point is not that the 109 never degraded aerodynamically during it`s development and when getting new stuff installed. The point is that it was later ALWAYS improved again. The Spit didn`t, as being said the ONLY aerodynamic improvement was the retractable tailwheel, that managed to get a few mph extra, and was used only on very few planes...
As for the 109s, it got two major aerodynamic facelift during it`s development, once in 1940 with the 109F, and in 1943 with the 109K. Which would explain you why it turned out that while in 1940 the Spitfire I was as fast with somewhat less HP, in 1944 it was 60 km/h slower with the same HP...
And just some comment on those drag factors... bulges for HMGs, yes, 9 km/h lost, which wouldn`t matter much in air combat, but the plane was now carrying effective MGs, instead of just ballast like the Spit did until 1944 with those .303s. And then in 1944 those HMG bulges gone.. The Spit had nice big bulges for the cannons through it`s service, THANK YOU desingers who designed a thin wing so much unsuited for bigger weapons, THANK YOU weapon developers in Britain who couldn`t came up with a cannon of their own, and forced the designers to put an large French cannon designed for rigid engine mounts and not flexible wings which cause them to jam.
bulges for the wheels, of course they had it, from 1943, the Spitty had them too, from 1941.
"numerous other little bulges and scoops", yeah, sure they existed somewhere if you say it. :D an air intake sticking out from the side of the fuselage
"an air intake sticking out from the side of the fuselage" - that`s the best one Milo, you are actually critisizing a better aerodynamic solution that reduced the drag.. :rofl Yep, sticking out the air intake was putting it out of the airstream, which meant an undistrupted boundery layer around the fuselage, the same reason the Mustang did the same with it`s radiator (perhaps because it`s designer was also a former Mtt engineer..)
3 radiators, WOW, Milo, getting desperate, huh? :D It ain`t the number of them counting, it`s the shape. The 109`s was sunken into the wings, having some 40mm sticking out in the front, the whole thing heavily profiting from the meredith effect.. On the Spit, it`s at least half a meter, I STILL don`t get who was that idiot who could design such a monstrosity onto a fighter.. literally drag bags they are, and aerodynamic test show they have serious turbulance inside as a bonus. Again the problem is not that.. the problem is it just worser and worser during the development!
ADF loop, yeah. That`s a classic. Huge drag I guess. :D
"a retractable tail wheel that was sometimes locked down," - hmm, yeah. I guess you mean that whereas the 109 had their retractable t/w *sometimes* locked down, the Spit never really had this problem, as it couldn`t retract the tail wheel in the first place. I wonder though what effect it had on the Spit`s tailwheel drag that it wasn`t lockable either.. lots of turbulance there again.
Gondola guns hung from the wings, yep, optional, raising firepower to a level a Spitfire never had, at a horrific loss of 5 mph. Now for comparison tropicalized Spits lost some 10-15 mph, and that was quite common. Their air intakes were on the nose chin, being as big as the oil cooler on the 109, nicely createing a lot of turbulance all the way behind, smarty positioned at the spot they can eat the most dust from the ground, which meant they needed a filter constantly used which added even more drag, and even then it was useless, the Soviets found the Spit IXs so sensible to dust that they only operated them from the well built airstrips of the PVO in the rear areas..
Mitchel put that plane very nicely together in the mid-1930s... very nice clean fuselage, good for it`s time, nice features, nice characteristics.. but from there it was a downhill run, the whole development can be described with two words : ad hoc.
-
The subject of this thread is the efficiency factor of the airplanes and how to determine it. I'm not comparing planes here but looking for what ballpark the e factor really was.
As an example the F2A-3 (http://members.tripod.de/luftwaffe1/aircraft/navy/F2A3_turning.pdf) had tested efficiency factor about 0,77.
If somebody has relevant information, opinion, or what so ever under this topic; please post! Otherwise cut the crap or start a new thread.
gripen
-
Good Gripen, good.
I have to have a potshot at Izzy though, forgive me please ;)
Firstly:
"The Spit had nice big bulges for the cannons through it`s service, THANK YOU desingers who designed a thin wing so much unsuited for bigger weapons"
So, it would actually have been better in your your opinion to have the wing much thicker?
Secondly:
"As for the 109s, it got two major aerodynamic facelift during it`s development, once in 1940 with the 109F, and in 1943 with the 109K. Which would explain you why it turned out that while in 1940 the Spitfire I was as fast with somewhat less HP, in 1944 it was 60 km/h slower with the same HP... "
Firstly, in 1940 the Spitfire MkI had roughly the same Hp on 87 oct, and somewhat more on 100 oct.
The Spitfire had less power at high alt, where max speeds were obtained.
Once fitted with a rotol 3-blade airscrew it outclimbed the 109 while being heavier, it was however a tad slower.
Once up to 100 oct the Spit I outclimbed and outran the 109E.
In 1944 you had Spit XIV swarming around. I have not seen data of a 109 being 40 mph faster. So please promote this.
Then here:
"THANK YOU weapon developers in Britain who couldn`t came up with a cannon of their own, and forced the designers to put an large French cannon designed for rigid engine mounts and not flexible wings which cause them to jam. "
You may have a problem here, for the Hispanos were very good weapons indeed. I have not stumbled across any tales of horrible jammings yet. Please promote som.
Very good ballistics, good ROF, and high velocity.
In 1941 perhaps the finest aircraft cannon in the world?
Then finally here:
"better aerodynamic solution that reduced the drag.. Yep, sticking out the air intake was putting it out of the airstream, which meant an undistrupted boundery layer around the fuselage, the same reason the Mustang did the same with it`s radiator (perhaps because it`s designer was also a former Mtt engineer..) "
You must have forgotten take your pills....
The radiator needs air to function, it always causes parasite drag.
Put it in the "shade" and you'll need it to be bigger.
Now, the boundary layer around the fuselage is not as important as say on the top of the wings. Do you know what a boundary layer is? A couple of days ago you did not know what a washout is.
You may of course have a problem with the Mustang, for on the same power as a 109 it was faster, with very much more range. Quite some headache for a brownie-trousers bubchen. But so be it, a German had his hand in on the design so it stays as a so-so.
Frankly I'd like to know what the heck is wrong with you. The only explanation I can think off is complete 1930's German religion, or a brownskirt reincarnation.
I mean, pointing at brilliant design features of objects like the 109, 262 etc is very fine. But you seem so obsessed with hatred on anything WW2 allied sided that it is just stunning. Be it tanks, armies, aircraft, campaigns, victories, political figures, the total outcome or whatever. Anything German just must be better.
Can you tell me why?
All ears?
Regards
Angus
p.s. Sorry gripen
:rolleyes:
-
Well, the problem is you have claimed things which are not true several times in this thread.
Bull Malarky
Your the one who keeps manipulating your data and using wrong figures to produce your results. You want us to compare a 1/6 wooden model Spitfire to the actual FW-190A8!!
You cherry pick the best performance data for the Spitfire while using the worst data for the FW-190. Data which even the flight testers admit they are not getting close to the extimated performance of the Horsepower output. Data where the flight testers cannot get the engine to idle without fouling plugs and the engine simply quits at altitude to the point the test cannot be completed.
Not the topic of this thread and learn German...
Your the one who needs to check your German. Here you attack me and in the other Thread ADMIT it does not say polished or filled!! It says primed and painted.
Obviously your like a small child having a tantrum when shown to be wrong.
Crumpp
-
Originally posted by Crumpp
Your the one who keeps manipulating...
Well, so far you have not posted anything relevant on efficiency factor nor anything which proves that model testing is bad nor proof for large scale wind tunnel in the Germany...
Why don't you just make your life easier and post relevant data or shut up.
gripen
-
nor anything which proves that model testing is bad nor proof for large scale wind tunnel in the Germany...
I don't what to post on that Gripen. Your a smart guy and have access to some aerodynamic information resources.
Why don't you prove it's not true?
I have told you to make a simple phone call to HTC and ask.
Crumpp
-
Originally posted by Crumpp
Why don't you prove it's not true?
Generally if somebody makes a statement, it's up to him/her to prove it true. The others do not need to disprove it.
If you want to continue, stay on topic and post relevant data if you can.
gripen
-
Sorry gripen but when Barbi goes on another one of his anti-Spit tirades I can't let it pass.
As can be seen in his reply, he is clueless of what he is babbling about in his delirious state.
-
If you want to continue, stay on topic and post relevant data if you can.
Exactly,
Your data from a 1/6 wooden model of the Spitfire is not relevant. You cannot make any meaningful direct comparision between data from a wooden model and compare it with data derived from an actual aircraft of a different type.
Now models are useful for determining general characteristics of the same aircraft. A wooden Spitfire Model is useful for getting a good idea of how an actual Spitfire will perform but not specifics.
You certainly cannot compare a 1/6th scale Wooden Spitfire to a real FW-190. You can only make very very General comparisions. Certainly not to the level of accuracy we are discussing.
This whole thread would have been halted from the begining had I know you were throwing up data off a wooden model. I just assumed you were smarter than that.
Unless you have a drag polar derived from an actual aircraft then you can only make comparisions using general formulas.
If you get this data please let me know.
Crumpp
-
Originally posted by Angus
Good Gripen, good.
I have to have a potshot at Izzy though, forgive me please ;)
[/B]
http://www.hitechcreations.com/forums/showthread.php?s=&threadid=136037
-
Originally posted by Crumpp
Exactly, blaah blaah...
I don't see a reason why you post to this thread If you can't bring in relevant data or you can't prove your arguments.
gripen
-
Your data from a 1/6 wooden model of the Spitfire is not relevant. You cannot make any meaningful direct comparision between data from a wooden model and compare it with data derived from an actual aircraft of a different type.
Crumpp
-
Well, only you are making comparisons in this thread and please point out the reason why the data from the wind tunnel with a model should not be comparable.
Besides the Fw data is probably determined with the models as well.
gripen
-
Here is the Cd/Cl^2 chart for the Spiteful. The data comes from the RAE wind tunnel tests with the 1/5,75 model:
(http://personal.inet.fi/koti/soon.moro/Spiteful.jpg)
This is a particularly interesting case because there was no washout and the resulting e factor is 0,81.
gripen
-
Gripen,
What does this:
http://1000aircraftphotos.com/APS/2918.htm
Have to do with anything?
Nice info but what does it tell us about the FW-190 vs Spitfire drag picture?
Crumpp
-
This thread is about determining efficiency factor of the WWII fighters. While the Spiteful did not saw service during war, it was designed during war and it's particularly interesting due to its wing which had no washout. And we are not directly talking about the absolute amount of drag but how the drag changes with the Cl.
This thread is not another X vs Y comparison, start a new thread if you are not interesting about determining efficiency factor.
gripen
-
From the back of my head, I rememberthe Spiteful being a bit of a disappointment. The wing was just marginally more effective than the Spitfire wing. ?!?!
-
Yep, there was just marginal improvement in the drag rise due to compressibility if compared to the original Spitfire wing.
gripen
-
This thread is not another X vs Y comparison, start a new thread if you are not interesting about determining efficiency factor.
Actually this thread was about if the general formulas using AR were viable to determine e factor. It was shown they are a good way to determine it. Having a drag polar from the actual aircraft is better.
And you cannot compare Wooden Models to actual aircraft of a different design.
Crumpp
-
Originally posted by Charge
"There were modifications withing the Mk IX series, and then there was the superb Mk VIII....."
The VII was stretched a bit and was mounted with 61, 64 or 71 Merlin and was the first to have those two similar radiators under the wings and it had a retractable tail wheel. The VIII was unpressurized variant with Vokes filter with a few different engines.
The tail wheel was however made fixed in IX and the late IX model had again the retractable one where as the other later models again had fixed wheel and again from XII onwards the tailwheel was retractable. Maybe just a feature of parallel development, dunno.
-C+
Basically it came down to the IX being developed off the Spit Vc as more of a lash up of the Merlin 60 series to get it into service faster hence the fixed tail wheel.
The VIII incorperated many of the refinements of the prototype Mk III that never went into production but had a retractable tail wheel, short span ailerons and at one point clipped wings.
The XII had roughly 50 of the hundred built on Spit Vc airframes so they had fixed tail wheels. These being the XIIs in the EN serial ranges, and about 50 of them built on Mk VIII airframes in the MB serial ranges which had retractable tail wheels.
And as the IX and VIII in many ways had similar performances, from a production stand point it was smarter to continue to improve on the IX and then XVI airframes without the refinements like retractable tail wheels and short span ailerons.
While the Spit VIII is considered the best overall of the Merlin Spitfires, the IX clearly was the most significant Merlin variant along with XVI, which was a Spit IX with an American built Packard Merlin 266 in place of the Rolls Royce built Merlin 66 in the IX
Dan/Slack
Late to the party as usual
-
Originally posted by Crumpp
Actually this thread was about if the general formulas using AR were viable to determine e factor. It was shown they are a good way to determine it.
Could you please show us where it was shown out that general formulas are good way to determine it?
The data shows that these gives almost routinely around 10% too high values for WWII fighters.
So far it has been also shown that a generalized value of e around 0,75-0,8 gives much better estimate than generalized formulas.
Originally posted by Crumpp
And you cannot compare Wooden Models to actual aircraft of a different design.
As noted above: If you make an argument you should be able to prove it and the purpose of this thread is not to make comparison.
gripen
-
Could you please show us where it was shown out that general formulas are good way to determine it?
Read Badboys comments on the General Formulas:
The number normally represented by the character e in induced drag calculations was originally known as Oswald’s efficiency factor, and his original paper is available for download from the NACA report server. More commonly it has a component of parasite drag lumped in with it and is just called the airplane efficiency factor and can be estimated depending on the aspect ratio, taper ratio, sweep angle and twist. Theoretically an elliptical wing would have an efficiency factor of 1, meaning that it will have a coefficient of induced drag close to the theoretical maximum. Even though it is a function of aspect ratio, sweepback angle, taper ratio and twist, the largest influence on the wing of a WWII fighter with very little sweep or twist comes from aspect ratio and taper ratio and so there are approximate formulae for estimating e that only include aspect and taper ratio, and even more approximate methods that only include aspect ratio. The important thing you must appreciate is that they are only approximate. But better than just assuming a constant value for every aircraft.
And certainly better than comparing a Wooden Model to an actual aircraft of a different design.
What they do show us is that Lednicer was correct. The Spitfire does have a slight advantage. The general conclusion you can draw off this thread is:
1. In high G manuvering and in the low speed realm at the induced drag "wall" the Spitfire has the advantage in the overall drag situation. The advantage in High Speed manuvering is only slight due to the FW-190's greater mass requiring greater braking forces.
2. Angus was correct when he concluded:
So, it's not an ALWAYS, but MOST OFTEN subject.
The FW-190 "MOST OFTEN" has less drag than the Spitfire.
It always has less form drag.
Crumpp
-
Originally posted by Badboy
The important thing you must appreciate is that they are only approximate. But better than just assuming a constant value for every aircraft.
The idea behind the generalized formula is good if we assume tapered wing with near optimal taper ratio. Apparently Wood's and NADC formula are created with somekind of statistics on the measured e factors. The problem is that that the WWII fighters generally had lower e factor which means that the both formulas give too high estimates and therefore also a constant value works better in most cases. As an example here is comparison with constant value of e 0,8:
(http://personal.inet.fi/koti/soon.moro/etest.jpg)
It can be seen that generalized formulas result about twice larger average error than the constant value. It can be also seen that there is a correlation between the aspect ratio and the e factor, so it is possible to create a better working general formula than Wood's or NADC.
Originally posted by Crumpp
And certainly better than comparing a Wooden Model... blaah blaah
Only you are comparing planes in this thread and please point out what is wrong with model tests. The lift distribution should be similar.
No proof, no argument.
gripen
-
It can be seen that generalized formulas result about twice larger average error than the constant value.
Gripen assuming a constant value for e in all aircraft is very much like trying to compare cars as if they all had the same engine.
It is unworkable from a comparision standpoint. Garbage in Garbage out.
Margin of error, that's laughable. So in order to have a margin we have to know the correct answer to see how far off of it we are in our calculations. So tell us, Just what exactly is the "e" factor for the REAL Spitfire and the FW-190!! Not a wooden model to the real plane.....
As you have been pouting about, the whole purpose of this thread to figure out the exact "e" factor. Now your saying you already have a base to judge the "margin of error"!
But better than just assuming a constant value for every aircraft.
Badboy teaches aeronautics for a living. I will take his word over yours anyday. No matter how many "margin of error" charts you throw up.
Crumpp
-
Originally posted by Crumpp
assuming a constant value for e ... blaah blaah
The reason for the constant value on the above chart is only the show what's the problem with above mentioned generalized formulas; in most cases both give too high value of e.
Originally posted by Crumpp
So tell us, Just what exactly is the "e" factor for the REAL Spitfire and the FW-190!!
My estimates are included above chart. Probably the service planes had even lower values.
gripen
-
My estimates are included above chart. Probably the service planes had even lower values.
So what is the exact value that you are using to determine this so called "Margin of Error"?
The reason for the constant value on the above chart is only the show what's the problem with above mentioned generalized formulas; in most cases both give too high value of e.
BS
There is no problem with the generalized formulas. You should listen to someone who has actually studied aeronautics instead of making up your own theories. Write NASA and tell them Wood and the NACD are wrong.
It's not the "exact" number that is important but it's place relative to other wings on the same scale. The "exact" number will change based on the formula and the scale used. Assuming a constant value creates a false picture of this position by totally ignoring the design of the wing.
Now how is that good? Except for the fact it artificially inflates the Spitfires standing on the scale.
The more of these design parameters you can account for the better. This thread started out on the right track. Too bad it was derailed by your wooden model which you attempted to pass off as the real thing.
Crumpp
-
Originally posted by Crumpp
So what is the exact blaah blaah...
Those are the numbers I got. I have given the sources as well as the quide lines for calculations. So far I have not seen you posting anything which proves that there is something wrong with the numbers I got.
For others, here is the drag polars for the F2A-3:
(http://personal.inet.fi/koti/soon.moro/F2A.jpg)
The interesting part is that they tested the F2A-3 also with flaps and the values of e are:
Flaps in 0,77
Flaps 22 deg open 0,80
Flaps 66 deg open 0,95
gripen
-
Those are the numbers I got.
Exactly.
So YOUR calculations using a constant value become the center point in which the "Margin of Error" is calculated.
Why...because you don't know the exact value and not having a drag polar off an actual aircraft are unable to even approximate it with any accuracy.
That's pretty funny!
The interesting part is that they tested the F2A-3 also with flaps and the values of e are:
They tested the FW-190 with and without prop, various power settings, landing gear up and down, and flaps in both positions. It's several pages of polar plots and prop efficiency charts.
Crumpp
-
Originally posted by Crumpp
Blaah Blaah Blaah...
I don't see anything relevant in your post.
gripen
-
this is why...
(http://www.geocities.com/galemcraig/strtvrtx2.jpg)
-
Exactly,
He keeps chasing his tail. Does not want to admit his data is faulty on the Spitfire.
Even when a man who teaches this stuff for a living sets him straight!
Crumpp
-
The NACA report on the wind tunnel tests with a 1/2,75 scale model of the F4U-1 can be found from here (http://naca.larc.nasa.gov/reports/1942/naca-wr-l-440/). The drag polar results e factor value 0,77.
We have now eight values of e factor for tapered winged WWII fighters with varying value of aspect ratio; not a large set but we can easily determine formula for e factor with these and compare results to the generalized formulas:
e = 0,9815 * AR ^(-0,1341)
In the table below the new formula is marked as e-f:
(http://personal.inet.fi/koti/soon.moro/ef.jpg)
As we can see the new formula gives much more accurate results, the error being one eight or less than with the generalized formulas. The next step would be adding taper ratio as well as washout to the formula but even in this form it works pretty well.
gripen
-
Originally posted by gripen
As we can see the new formula gives much more accurate results, the error being one eight or less than with the generalized formulas.
gripen
Ok, that seems to be about it for this thread, so let’s see where it has lead us.
1) 1933 Oswald publishes his original paper and it is available for download from the NACA report server. Glauert was then first to derive a theoretical expression for rectangular wings, in the form of a correction to the value for elliptical wings that could be applied to Oswald’s work.
2) 1935 Professor Wood published his empirical methods, which are based on flight test data.
3) 1949 Perkins & Hage publish Oswald’s methods along with other eminent aerodynamicists of the time, including Dwinnel & Jones. Infact, every aerodynamics text book since includes and supports that work.
4) Dr Raymer publishes another empirical method based on NADC data from 1966 in his book on aircraft design in 1999.
5) Lednicer produces similar results using computational aerodynamics and publishes his results in an article in 1999.
6) Then in 2004, Gripen claims they are all wrong and that his method is better. He publishes a new concept, the e value for every plane is practically the same and entirely refutes thousands or hours of research and negates the work of some truly great men…
Gripen… you have not a shred of credibility, give it up already. LOL!!
Dweeb
-
Originally posted by Dweeb
1) 1933 Oswald publishes his original paper and it is available for download from the NACA report server. Glauert was then first to derive a theoretical expression for rectangular wings, in the form of a correction to the value for elliptical wings that could be applied to Oswald’s work.
Hm... Actually it was von Prandtl (1918) who introduced whole theory of the lift distribution. Some time later (1926) Glauert published his formulas for the rectangular wings, it should be noted that his formulas as well as Glauert's are for the wing only. Oswald published his work 1933 and he introduced the term of the e factor for the entire airframe.
Originally posted by Dweeb
2) 1935 Professor Wood published his empirical methods, which are based on flight test data.
I don't know what kind of data Wood used for his empirical analysis. But we can be sure that there is no any WWII fighters.
Originally posted by Dweeb
3) 1949 Perkins & Hage publish Oswald’s methods along with other eminent aerodynamicists of the time, including Dwinnel & Jones. Infact, every aerodynamics text book since includes and supports that work.
Yep, and these are the methods I have used for my analysis.
Originally posted by Dweeb
4) Dr Raymer publishes another empirical method based on NADC data from 1966 in his book on aircraft design in 1999.
Again, I don't know what kind of data set is used for the NADC dataset but it's probably based on something else than WWII fighters.
Originally posted by Dweeb
5) Lednicer produces similar results using computational aerodynamics and publishes his results in an article in 1999.
Lednicer calculated lift distribution for three planes and did not claim any specific value for the e factor.
Originally posted by Dweeb
6) Then in 2004, Gripen claims they are all wrong and that his method is better. He publishes a new concept, the e value for every plane is practically the same and entirely refutes thousands or hours of research and negates the work of some truly great men…
Thank you for great logic, seems that you best even Crumpp in that area.
What I really say above is that generalized formulas by Wood and NADC seem to give too high e factor for the WWII fighters and this is backed up by empirical data. The most probable reason for this is that these formulas are probably based on something else than WWII fighters. Most WWII fighters had very similar dimensions and therefore also the values of the e factor are very similar.
gripen
-
Originally posted by gripen
Hm... Actually it was von Prandtl (1918) who introduced whole theory of the lift distribution. Some time later (1926) Glauert published his formulas for the rectangular wings, it should be noted that his formulas as well as Glauert's are for the wing only. Oswald published his work 1933 and he introduced the term of the e factor for the entire airframe.
gripen
Wrong, Oswald was first, that's why they named it after him! LOL!
Gripen, I'm sorry but this has to be said... Those experts and their data are all correct... You are just a confused amateur with a deplorable attitude and no credibility.
The data you have, has been available to aerodynamics experts for decades, and during analysis of WWII fighters throughout the 40's they all applied it, because they all acknowledge Oswald's theory.
Even today modern universities use those methods, and still teach it, it appears in almost every text book on the subject.
If you have discovered that they are all wrong, and that the e value for all WWII fighters was really the same, point us to your learned papers or journal publication... LOL!
You started this thread confused, and then declined from there! It is time for you to stop wasting everyone's time with your nonsense... let it go!
Dweeb
-
Originally posted by Dweeb
Wrong, Oswald was first, that's why they named it after him! LOL!
Why don't you just read what Oswald really wrote (http://naca.larc.nasa.gov/reports/1933/naca-report-408/index.cgi?page0002.gif); he directly refers von Prandtl's theory.
Glauert published his works actually 1926/1929 and these are used as sources (http://naca.larc.nasa.gov/reports/1933/naca-tn-483/index.cgi?page0019.gif) in many NACA documents as well as other publications.
What Oswald did was just continuation of the work by von Prandtl, Glauert and others, while Prandtl and Glauert handled lift distribution along the wing as relative factor, Oswald expressed theory to the drag of the entire airframe. Besides Oswald's report is actually about performance calculations and the efficiency factor is just one small but important part of it.
Originally posted by Dweeb
Gripen, I'm sorry but this has to be said... Those experts and their data are all correct... You are just a confused amateur with a deplorable attitude and no credibility.
Of the mentioned persons only Perkins&Hage and Lednicer actually have studied WWII fighters and what I have calculated above is strictly based on their works. And the typical e factors claimed by Perkins&Hage are 0,7-0,85 (p. 95).
Originally posted by Dweeb
If you have discovered that they are all wrong, and that the e value for all WWII fighters was really the same, point us to your learned papers or journal publication... LOL!
Could you please point out from where you have got the idea that I have claimed that mentioned persons are all wrong?
All I say above is that generalized formulas by Wood and NADC seem to give too high value of the e factor for WWII fighters and this conclusion is backed up by empirical data. The most probable reason for this is that these formulas are created for different kind of airplanes; say relative large winged and slow planes combined with small fuselage. According to your logic this means that I claim all written about lift distribution and e factor wrong.
Generally if you want to discuss, please use logical and verifyable arguments.
gripen
-
Originally posted by gripen
Generally if you want to discuss, please use logical and verifyable arguments.
gripen
That’s very funny, coming from a guy who is quoting from a book that he doesn’t understand (or you wouldn’t have opened this thread with such a dumb question), can’t even spell the word verifiable, and uses third party wind tunnel data to produce completely erroneous results. Even better, after starting this thread without a clue, you already claim to have revolutionized the theory. Don't you think it is time you faced the fact that you don't have the knowledge, the experience, or the credentials to be carrying out any kind of aerodynamics research, let alone refuting the very solid work of renowned experts, that has been accepted for decades... You are not just a bumbling amateur, you are a fool who unsuccessfully tries to sound as though you know something by flooding threads with so much relentless nonsense, that everyone else gives up.
If you want anyone to believe your silly claims, do some original research, have your academic work published and then point us to the publication... The simple fact is, even a first year aero major would laugh at your nonsense, so why don't you just give it up and stop wasting everyone’s time with your complete and utter baloney.
Dweeb
PS Happy Thanksgiving!
-
Dweeb,
I have claimed all my sources which include wind tunnel data from various sources as well as flight tested data. I have given the guidelines for the calculations as well as the results. If my errors are so easy to see and laughable, why don't you just point them out.
gripen
-
Originally posted by gripen
Dweeb,
I have claimed all my sources which include wind tunnel data from various sources as well as flight tested data. I have given the guidelines for the calculations as well as the results. If my errors are so easy to see and laughable, why don't you just point them out.
gripen
Gripen,
I don't need to, because you don’t have an ounce of credibility. You claim to have found that Oswald’s efficiency factor is practically constant for a wide range of fighter aircraft built during the war years. If it were true it would mean that people like Oswald, Professor Wood and every other professional in the field has been mistaken all this time, and that Oswald, Professor Wood and Dr Raymer have wasted their time, wasted fortunes in research, and years of effort only to produce results that while being accepted in aerodynamics for decades, you have easily refuted it with some third party data. You claim the number was a constant after all, but not only that, you further claim to have found out what it is… As I said, if you want to be taken seriously with such laughable nonsense, do the same thing as those before you, do your research, present your paper for peer review and then publish! You will then either become famous or expose yourself as a fool. I know which one my money is on… LOL!
Dweeb
-
Originally posted by Dweeb
I don't need to, because you don’t have an ounce of credibility.
Sounds that you are not discussing here, just blaming me for something without proof.
Originally posted by Dweeb
You claim to have found that Oswald’s efficiency factor is practically constant for a wide range of fighter aircraft built during the war years.
Why the efficiency factor should not be similar among the planes with very similar dimensions?
Besides the data for the Bf 109G, Fw products, F4U and F2A is above. Please point out my errors.
gripen
-
In another thread (http://www.hitechcreations.com/forums/showthread.php?s=&threadid=136469) rose a question about the drag of the P-80, NACA data from the flight tests and wind tunnel test with a 1/3 model can be found from here (http://naca.larc.nasa.gov/reports/1948/naca-rm-a7i16/naca-rm-a7i16.pdf).
The e factor of the P-80 seems to be about 0,75 at Cl range 0,2- 0,8. Which again fits very well to the other data seen in this thread.
Regarding accuracy of the wind tunnel data, this report gives also interesting read (http://naca.larc.nasa.gov/reports/1948/naca-rm-a7i16/index.cgi?page0008.gif):
"The agreement of the flight and wind-tunnel data is excellent at all Mach numbers of the test"
Dweeb is again most wellcome to show my errors.
gripen
-
"The agreement of the flight and wind-tunnel data is excellent at all Mach numbers of the test"
Again you need an english lesson Gripen.
The report says in the very NEXT sentence that the results in the low speed realm where "fortuitous" .
http://dictionary.reference.com/search?q=fortuitous&r=67
Happening by accident or chance. See Synonyms at accidental.
Crumpp
-
Sweet Jeezus.
You are incurable hagglers.
Now please be nice :D
-
Originally posted by Crumpp
The report says in the very NEXT sentence that the results in the low speed realm where "fortuitous" .
What it exactly says is:
"The close agreement between the low-speed data may be partly fortuitous considering that flight-test was computed from the thrust (the predominate force at low speed) taken from an engine calibration chart."
This thread is about e factor and lift distribution not semantics.
gripen
-
Originally posted by gripen
Besides the data for the Bf 109G, Fw products, F4U and F2A is above. Please point out my errors.
Why? So you can get a free education?
The simple fact is, you have already asked some pretty dumb questions, like the one you opened the thread with, and had some excellent answers by some very knowledgeable people, who according to you are all wrong. Despite the help you have received already, you still claim that the books that explain the theory, even the guy who has had it named after him, and Professor Wood, and Dr Raymer are all wrong, because you claim to have found that the airplane efficiency factor, is not a factor after all, it is a constant! Can’t you see how absurd that is? Do you really think that after decades of work, and countless years of research, it was all wasted because you are the first person to figure it out? Yeah right!
Dweeb
-
Originally posted by gripen
Why the efficiency factor should not be similar among the planes with very similar dimensions?
LOL, you might as well ask why any of the aerodynamic coefficients “should not be similar among planes with very similar dimensions?” but nobody is suggesting they all have the same lift coefficient, or the same drag coefficient, why then do you suppose they have the same efficiency factor? Please prove your argument, as you so often like to say.
Dweeb
-
Originally posted by gripen
I don't see a reason why you post to this thread If you can't bring in relevant data or you can't prove your arguments.
gripen
Gripen, trust me you haven’t proven anything here yourself, despite your pompous claims! None of your endless diatribe constitutes proof, it lacks the transparency and rigour of proper academic exposition and actually qualifies as little more than the ranting of a spoilt child trying to gain influence in a gamers board.
Dweeb
-
Umm, Dweeb, not quite getting where you're going.
That is regarding this:
"--------------------------------------------------------------------------------
Originally posted by gripen
Why the efficiency factor should not be similar among the planes with very similar dimensions? "
I guess he means RELATIVE dimensions. Correctly scaled that is.
Please enlighten!
-
Originally posted by Angus
I guess he means RELATIVE dimensions. Correctly scaled that is.
Please enlighten!
Yes, Gripen appears to be saying that because the aircraft in question (basically all WWII aircraft) have similar dimensions, they should have similar efficiency factors. The idea is complete nonsense for a number of reasons. Firstly, we are talking about a variety of aircraft and most of their dimensions are not really that similar, despite some common tendencies. The fact that the Aircraft in question have wings with different airfoil sections, means that they all have different maximum lift coefficients, different lift curve slopes, different drag coefficients and so on, does anyone believe those things should all be the same for every aircraft… of course not! And that’s not even including the variety of three dimensional factors like wing taper, sweep and twist, all things that have an influence on the efficiency factor, and when you consider the variety of WWII aircraft, there is no reason to believe the efficiency factor should all be the same, anymore than you would believe they all have the same lift coefficient. And even then we have not included every factor that can influence the results… Yet, Gripen is still happy to claim that they are all dimensionally similar, when they really aren’t, and to claim that is the reason they should all have a constant efficiency factor, which of course they don’t. Nobody believes they do accept him! It just gets more and more laughable every time he posts!
If you want to judge for yourself, just pick up some of the publications mentioned in this thread and realise that if the efficiency factor was constant, they would all be wrong, decades of well established aerodynamics, all down the drain because Gripen has discovered that the efficiency factor is really a constant… The man started with no credibility and has just gone downhill from there.
Dweeb
-
Originally posted by Dweeb
Why? So you can get a free education?
Determining the e factor from the drag polar or from the value of the K is quite simple as pointed out out in the Perkins& Hage:
(http://personal.inet.fi/koti/soon.moro/ERREE.JPG)
If you want to discuss, please point out my errors.
Originally posted by Dweeb
The simple fact is, you have already asked some pretty dumb questions, like the one you opened the thread with, and had some excellent answers by some very knowledgeable people, who according to you are all wrong.
The simple fact is that without several assumptions as with the generalized formulas, there is no way to determine e factor from the aspect ratio only.
Originally posted by Dweeb
Despite the help you have received already, you still claim that the books that explain the theory, even the guy who has had it named after him, and Professor Wood, and Dr Raymer are all wrong, because you claim to have found that the airplane efficiency factor, is not a factor after all, it is a constant!
Please show where I have claimed that Oswald, Wood and Raymer are all wrong. AFAIK they never analyzed WWII fighters but Perkins&Hage and Lednicer actually did.
All I say above is that generalized formulas by Wood and NADC seem to give too high value of the e factor for WWII fighters and this conclusion is backed up by empirical data.
Originally posted by Dweeb
Yes, Gripen appears to be saying that because the aircraft in question (basically all WWII aircraft) have similar dimensions, they should have similar efficiency factors. The idea is complete nonsense for a number of reasons. Firstly, we are talking about a variety of aircraft and most of their dimensions are not really that similar, despite some common tendencies. The fact that the Aircraft in question have wings with different airfoil sections, means that they all have different maximum lift coefficients, different lift curve slopes, different drag coefficients and so on, does anyone believe those things should all be the same for every aircraft… of course not! And that’s not even including the variety of three dimensional factors like wing taper, sweep and twist, all things that have an influence on the efficiency factor, and when you consider the variety of WWII aircraft, there is no reason to believe the efficiency factor should all be the same, anymore than you would believe they all have the same lift coefficient. And even then we have not included every factor that can influence the results… Yet, Gripen is still happy to claim that they are all dimensionally similar, when they really aren’t, and to claim that is the reason they should all have a constant efficiency factor, which of course they don’t. Nobody believes they do accept him! It just gets more and more laughable every time he posts!
The data for the P-80, Bf 109G, Fw products, F4U and F2A is above. Please point out my errors. The Fw series is a good example because it contains three different aspect ratios.
gripen
-
"The close agreement between the low-speed data may be partly fortuitous considering that flight-test was computed from the thrust (the predominate force at low speed) taken from an engine calibration chart."
You honestly have no idea what that sentence is saying do you, Gripen??
Crumpp
-
Crumpp,
Let's look at the whole text:
"The agreement of the flight and wind-tunnel data is excellent at all Mach numbers of the test. The close agreement between the low-speed data may be partly fortuitous considering that flight-test was computed from the thrust (the predominate force at low speed) taken from an engine calibration chart. The drag data at high Mach numbers are on a better basis for comparison because the flight-test drag weas computed principally from from gravitaional components, jet thrust being of secondary importance. The flight results are taken from data previouysly issued in preliminary form. Refinements in calibration of the flight-test instruments have been made since the data were first issued."
Shortly it says that the agreement between the data sets is excellent, they had some doubts about low speed data (due to used thrust data) but the at high mach numbers data sets should be very comparable.
Besides it should be noted that used thrust values do not affect lift distribution.
gripen
-
Here is a chart which shows how the all data points mentioned above as well as above mentioned formulas fit in the AR/e chart, note that in this chart all the datapoints are calculated using least square method to fit linear line (the only notable difference is the Mtt data for the Bf 109G, e being 0,69 with this method):
(http://personal.inet.fi/koti/soon.moro/comp.jpg)
It can be easily seen that in the case of the tapered winged WWII fighters generalized formulas by Wood and NADC give allways too high value of the e. Only in the case of the wind tunnel data for the Spitfire (elliptical winged) these give about right value.
Another thing which can be easily seen is that aspect ratio explains poorly differences in the e factor regardless the used formula. Including the Spitfire, the value of the R^2 is about 0,2 in the best case. However in the case of the tapered winged planes (exluding Mtt data), the value of the R^2 is about 0,5 using the e-f formula (which can be easily altered further if more relevant data comes in).
gripen
-
Text says:
"The agreement of the flight and wind-tunnel data is excellent at all Mach numbers of the test. The close agreement between the low-speed data may be partly fortuitous considering that flight-test was computed from the thrust (the predominate force at low speed) taken from an engine calibration chart.
Gripen says:
Shortly it says that the agreement between the data sets is excellent, they had some doubts about low speed data (due to used thrust data) but the at high mach numbers data sets should be very comparable.
And attempts to pass off this off as proof that wooden models can be compared to actual aircraft of a different design here:
Regarding accuracy of the wind tunnel data, this report gives also interesting read:
"The agreement of the flight and wind-tunnel data is excellent at all Mach numbers of the test"
As for your repeated request to show your math errors:
Please point out my errors.
That would be a complete waste of time. It is not a math error your making but a conceptual one. Showing any error committed on your part is like beating your head against the wall. Look how many times the "waxed and polished" came up. You still contest the Focke Wulf performance data especially when it comes from FOCKE WULF! You would rather take the allies even when it proven the data is bad.
Crumpp
-
Originally posted by Crumpp
And attempts to pass off this off as proof that wooden models can be compared to actual aircraft of a different design here:
The NACA report compares the model and the real aircraft of the same design, and found out that the agreement was excellent.
Originally posted by Crumpp
As for your repeated request blaah blaah
Well, my calculations as well as used data are above for all who really want to discuss as well as for those who don't want to discuss (namely you and Dweep). In the case of the Dweep this apparently means that e factor is some how forbidden subject and bringing in data on this is same as claiming everything written on e factor not true.
gripen
-
The NACA report compares the model and the real aircraft of the same design, and found out that the agreement was excellent.
So this is a basis for saying all wooden models can now be directly compared with actual aircraft of a different design?
Well, my calculations as well as used data are above for all who really want to discuss as well as for those who don't want to discuss (namely you and Dweep).
What part of your concepts are flawed and not your math do you not pick up on?
Crumpp
-
Originally posted by Crumpp
So this is a basis for saying all wooden models can now be directly compared with actual aircraft of a different design?
The lift distribution and the form drag should be about the same regardsless the size of the model in the wind tunnel regardsless differences in the viscous drag, given the corrections are right. So I don't see the reason why the e factor determined from the wind tunnel data should not be comparable with another wind tunnel data.
Besides as noted above, the Fw data is probably determined with the models.
Originally posted by Crumpp
What part of your concepts are flawed and not your math do you not pick up on?
The data is above, feel free to play with it.
gripen
-
The lift distribution and the form drag should be about the same regardsless the size of the model in the wind tunnel regardslessdifferences in the viscous drag, given the corrections are right. So I don't see the reason why the e factor determined from the wind tunnel data should not be comparable with another wind tunnel data.
Besides as noted above, the Fw data is probably determined with the models.
And the wheel goes round and round and round and round.....
You got a copy of the full report?
The data is above, feel free to play with it.
YOUR data is above. Why "play with it" when it is based on flawed science. Garbage in Garbage out.
Crumpp
-
Tell me Gripen,
What are the differences between the Mach and Low Speed Realm of flight?
Crumpp
-
Originally posted by Crumpp
And the wheel goes round and round and round and round.....
Well, you are most wellcome to post relevant data. I have seen pictures of a Fw 190 in the Chalais-Meudon but not pictures of Ta 152s in the large scale tunnel. Besides all WWII era large scale tunnels were low speed tunnels, maximum air flow around 50-60 m/s.
gripen
-
Originally posted by Crumpp
What are the differences between the Mach and Low Speed Realm of flight?
See the drag data in the P-80 report.
gripen
-
Well, you are most wellcome to post relevant data.
That is exactly what we are waiting on from you. Some relevant data.
Such as:
Lets see and compare total drag @ 440fps at the Spitfires FTH altitude with the correct data:
BF 274 - 795.641828 @ 440 fps
Total Drag FW-190A5 -718.2845242@ 440 fps
BS354 - No data listed except the A/C was 7 mph slower w/50 bhp more output than the Merlin 66 @ (+18)
BS543 - 810.863789
FW-190A5 - 705.7484202
BS551 - 758.4823884
FW-190A5 - 704.890256
EN524 with 4 bladed prop - 765.4635575
FW-190A5 - 702.4712376
BS310 with 4-bladed prop - 793.0539968
FW-190A5 - 751.8834689
BS310 with 5 bladed prop - 794.0663191
FW-190A5 - 749.7161496
JL165 - done to death
MA 648 Merlin 66 (+18) SU Pump - 838.0549776
FW-190A5 - 804.8037275
BS310 - 792.6552588 Parasitic drag - 691.672177
FW-190A8 - 794.3964059 Parasitic drag - 606.4158622
At 315mph the total drag situation changes completely in the FW-190A8 favor.
JL165 vs FW-190A8 has been done to death, again in the FW-190A8's favor.
MA 648 is the ONLY Spitfire Mk IX to beat the FW-190A8 for drag and here is why:
quote:
--------------------------------------------------------------------------------
External equipment of the above four aircraft was similar with the exception that MA.648 had the new pattern of air intake.
--------------------------------------------------------------------------------
MA648 - 837.1701001 parasitic drag - 751.6212076
FW-190A8 - 869.2121249 parasitic drag - 709.9624067
The parasitic drag is always in the FW-190's favor but the total drag drops to within a few pounds of each other at 315 Mph but never swings in the FW-190's favor.
So ONE example of the Spitfire Mk IX with an experimental air intake beats the FW-190A8 in total drag. The others are behind their FW-190A contemporary the whole way.
That about covers the Spitfire Mk IX's vs. their contemporary FW-190 adversary.
Now this works out the same NO MATTER WHOSE formula you use for "e" factor. At least all the normal accepted ones. It probably won't once Gripen finishes his "theory".
Crumpp
-
Originally posted by Crumpp
Blaah Blaah...
Not relevant data nor nothing on subject of this thread, this thread is about determining the e factor.
I quess you have posted more than 100 times to this thread and so far nothing relevant.
gripen
-
Not relevant data nor nothing on subject of this thread, this thread is about determining the e factor.
That data is extremely relevant. No matter what accepted formula or theory you use the conclusions remain the same.
The parasitic drag is always in the FW-190's favor but the total drag drops to within a few pounds of each other at 315 Mph but never swings in the FW-190's favor.
The whole point is Gripen:
When the current proven theories do not produce the results you want you simply make up your own theory!!!
Pathetic.
Crumpp
-
Originally posted by Pyro
2- Threads should remain on topic, do not "hijack" topics.
-
How is it off topic?
The whole point of this thread is:
this thread is about determining the e factor.
And according to ALL the accepted Aeronautical Theories this conclusion is true no matter what formula you use:
The parasitic drag is always in the FW-190's favor but the total drag drops to within a few pounds of each other at 315 Mph but never swings in the FW-190's favor.
You don't like those results and are inventing your own "efficiency factor according to Gripen" theory in order to produce results that are more palatable in your spitfire fandom.
It is very relevant to this thread when you are shown it is wrong. Problem is you want a fan club and folks to be impressed with your drivel. You can't accept being wrong. Look at the FW-190 finish argument and the ridiculous data you want to use. Of course you get results that are wrong. Garbage in, Garbage out.
Rather than say "Thanks I did not know that" and move on. You spend enormous effort defending the indefensible.
So chuck your ridiculous theory were all WWII fighters have a constant "e" factor and let's get back to examining other designs. You're wrong again on this one.
Crumpp
-
Originally posted by Crumpp
Blaah Blaah Blaah...
This thread is about determining the efficiency factor. The theoretical back round can be found from the Perkins&Hage and my findings from the chart below, the sources are given above.
(http://personal.inet.fi/koti/soon.moro/comp.jpg)
Bring in relevant data if you want to argue.
gripen
-
This thread is about determining the efficiency factor. The theoretical back round can be found from the Perkins&Hage and my findings from the chart below, the sources are given above.
Not for your theory. Please list the theoretical background again why NACA, Woods, AND Perkins&Hage are wrong. You only use the chart from Perkins&Hage, Gripen. They use Glauert's theories same as Woods and NACA. So don't claim you are using Perkin&Hage's theories. Only their chart to post YOUR CACULATIONS finding on.
Bring in relevant data if you want to argue.
That is exactly what others have tried to do. Unfortunately according to you only YOUR FINDINGS are classified as relevant data. What you want is "Gosh that is great!" ego stroke.
Pathetic
Crumpp
-
Originally posted by Crumpp
Blaah Blaah Blaah...
So far you have posted nothing relevant. The data is above, theory is in Perkins&Hage and my calculations are strictly based on these.
gripen
-
--------------------------------------------------------------------------------
This thread is about determining the efficiency factor. The theoretical back round can be found from the Perkins&Hage and my findings from the chart below, the sources are given above.
--------------------------------------------------------------------------------
So far you have posted nothing relevant. The data is above, theory is in Perkins&Hage and my calculations are strictly based on these.
Not for your theory. Please list the theoretical background again why NACA, Woods, AND Perkins&Hage are wrong. You only use the chart from Perkins&Hage, Gripen. They use Glauert's theories same as Woods and NACA. So don't claim you are using Perkin&Hage's theories. Only their chart to post YOUR CACULATIONS finding on.
quote:
--------------------------------------------------------------------------------
Bring in relevant data if you want to argue.
--------------------------------------------------------------------------------
That is exactly what others have tried to do. Unfortunately according to you only YOUR FINDINGS are classified as relevant data. What you want is "Gosh that is great!" ego stroke.
Pathetic
Crumpp
-
Originally posted by Crumpp
Blaah...
The data and the calculations are above, please point out errors.
gripen
-
The data and the calculations are above, please point out errors.
The calculations are listed nowhere in this thread. Only your results.
Please quit claiming you are using Perkins&Hage's theory. It is deceitful of you to do so. You are only using the chart from their book. They use the same theories you are trying to discredit.
Crumpp
-
It's simple, just form the Cl^2 chart from the drag polar and fit line to the linear part with your favorite system (solver or what ever) then just calculate e as shown above step by step. If this goes over your head, it's not my problem.
It's exactly the same system as in Perkins&Hage.
(http://personal.inet.fi/koti/soon.moro/ERREE.JPG)
gripen
-
Gripen,
Posting an extract from Perkins & Hage, only provides further evidence that the efficiency factor is indeed a factor and not a constant, if it were a constant, they would have called it that, and used it as a constant, but they don’t, they acknowledge it as it is, and every aero engineer since has accepted it as such. Nobody, since its conception has ever treated it as a constant… until you came along! LOL!
As for your wind tunnel data, Professor Wood says in his paper:
Originally stated by Professor Wood
Performance estimates based on model test results are unsatisfactory because of the large and uncertain corrections necessary to make the results correspond to the free flight of a full sized airplane.
He is referring to the fact that model tests involving fluid forces never completely represent the full scale condition because of practical limits on the control of the variables involved. For example, the density and viscosity of the fluid are controllable only within narrow limits depending on the range of pressures and temperatures within the test apparatus, and they can vary significantly. Most model tests involve errors relating to Reynolds number, Mach number surface irregularities and roughness, turbulence, and interference effects. Infact, before and during WWII the attempts to correct for differences in turbulence effects between wind tunnels and free air proved so inaccurate that full scale flight testing was widely considered to be essential for anything other than preliminary design. These considerations make the data you have, based on wind tunnel tests made on scale models during the period in questions, almost entirely useless, and lead to conclusions that are not only invalid, but in your case, actually fly in the face of accepted aerodynamic theory.
Hardly surprising, but since you have no formal aerodynamics training, you wouldn’t have been expected to know any of this, you probably only picked up a copy of Perkins and Hage when you saw Badboy’s post as quoted here:
Originally posted by Badboy
There are a large number of sources that clearly state the relationship between the efficiency factor and Aspect Ratio (AR) and several of them also provide graphs of e against AR, examples include:
Principles of Aerodynamics, by Dwinnell, page 144.
Airplane Performance Stability and Control, by Perkins & Hage, page 73.
Elements of Practical Aerodynamics, by Jones, page 73.
Some others publications have multiple curves in order to include taper ratio. A readily available source, for which you can also obtain software, is:
Aerodynamics Aeronautics and Flight Mechanics, by McCormick, page 172.
In this thread:
http://www.hitechcreations.com/forums/showthread.php?threadid=128143&referrerid=4117
Because from that point forward you appear to have clued up slightly, however, you still seem to be determined to rely on specious data. The quotes above from Professor Wood, explain why he used full scale flight tests to obtain his data. It explain why, in 1949 four years after WWII, Perkins & Hage were still calling Oswald’s discovery, an Efficiency Factor, not a Constant!
If you want anyone to take your wind tunnel data seriously, you need to post a lot more information about it, for example, the method of supporting the model and the control of associated turbulent flow problems, and the degree of detail in the model, for example were details such as the radiator details, intakes, pitot heads, radio masts, airspeed and air temperatute booms, antenna, guns or gun ports, gun and ammunition door and hatches included on the model? Were any other measures taken to simulate discontinuities in those locations? What corrections were made for Mach number and dynamic pressure calibration? What corrections were made for Reynold’s number? Was the static pressure survey in the tunnel conducted with multiple booms and how were local air disturbances handled? What corrections were made to the Mach number and aerodynamic coefficients for the blockage factor due to the model? What corrections were made to the angle of attack and drag coefficients due to the presence of the tunnel walls? Those are the sort of details you need to be discussing, along with some very self critical estimates of the size of the error bars involved. That was enough to cause Wood to say:
Originally stated by Professor Wood
Performance estimates based on model test results are unsatisfactory because of the large and uncertain corrections necessary to make the results correspond to the free flight of a full sized airplane.
When he conducted his tests, and concluded that e was not constant. Dr Raymer published his empirical method based on NADC data from 1966 in his book on aircraft design in 1999, and he still considered that e was not constant… The reason is that those men are professional aerodynamic engineers, with vast experience and high standing in their fields… You on the other hand are a game player using specious data to win an argument in a silly thread.
So, why don’t you stop pretending to know more than you do, instead of making grossly incorrect statements about a subject you are clearly struggling to pick up as you go along? You would make much better progress by asking genuine questions, instead of acting as though you actually already know what you are talking about, while making absurd claims. Your approach is more befitting a clown than someone with a genuine interest in the subject and a real desire to learn.
Please, give up your silly claims, or compile them with the proper transparency and rigour of a serious academic exposition and publish your findings… Put up, or shut up!
Dweeb
-
In fact Perkins&Hage Clearly states the relationship between Aspect Ratio and "e" Factor. It is a key element.
(http://www.onpoi.net/ah/pics/users/503_1101598168_phinduceddrag.jpg)
Ohh look, here they refer to Oswalds efficiency factor:
(http://www.onpoi.net/ah/pics/users/503_1101597886_phefactor.jpg)
How did they figure that out? With the same forumla used to make these calculations!
Lets see and compare total drag @ 440fps at the Spitfires FTH altitude with the correct data:
BF 274 - 795.641828 @ 440 fps
Total Drag FW-190A5 -718.2845242@ 440 fps
BS354 - No data listed except the A/C was 7 mph slower w/50 bhp more output than the Merlin 66 @ (+18)
BS543 - 810.863789
FW-190A5 - 705.7484202
BS551 - 758.4823884
FW-190A5 - 704.890256
EN524 with 4 bladed prop - 765.4635575
FW-190A5 - 702.4712376
BS310 with 4-bladed prop - 793.0539968
FW-190A5 - 751.8834689
BS310 with 5 bladed prop - 794.0663191
FW-190A5 - 749.7161496
JL165 - done to death
MA 648 Merlin 66 (+18) SU Pump - 838.0549776
FW-190A5 - 804.8037275
BS310 - 792.6552588 Parasitic drag - 691.672177
FW-190A8 - 794.3964059 Parasitic drag - 606.4158622
At 315mph the total drag situation changes completely in the FW-190A8 favor.
JL165 vs FW-190A8 has been done to death, again in the FW-190A8's favor.
MA 648 is the ONLY Spitfire Mk IX to beat the FW-190A8 for drag and here is why:
quote:
--------------------------------------------------------------------------------
External equipment of the above four aircraft was similar with the exception that MA.648 had the new pattern of air intake.
--------------------------------------------------------------------------------
MA648 - 837.1701001 parasitic drag - 751.6212076
FW-190A8 - 869.2121249 parasitic drag - 709.9624067
The parasitic drag is always in the FW-190's favor but the total drag drops to within a few pounds of each other at 315 Mph but never swings in the FW-190's favor.
So ONE example of the Spitfire Mk IX with an experimental air intake beats the FW-190A8 in total drag. The others are behind their FW-190A contemporary the whole way.
That about covers the Spitfire Mk IX's vs. their contemporary FW-190 adversary.
Now this works out the same NO MATTER WHOSE formula you use for "e" factor. At least all the normal accepted ones. It probably won't once Gripen finishes his "theory".
Crumpp
-
Dweeb,
Hm... When I started this thread, only thing I knew was that generalized formulas seem to give quite high value of the e. Since then I have studied quite a lot drag data and so far I can conclude that these generalized formulas really give too high value of the e.
What I have calculated above is based on tested data, in wind tunnel and also in the flight (models as well as real planes). I have no idea how Wood or NADC created their formulas but probably not using data from the WWII fighters.
I have not claimed e factor constant, Perkins&Hage makes this quite clear and this can be seen partially from the wind tunnel data as well as flight tested data. However, in many cases there is a linear stage as pointed out in the Perkins&Hage.
I have not claimed that the e factor is same for all WWII fighter planes, again data shows that there is variation but not particularly due to the aspect ratio.
The problem here is that you join in the discussion with the attitude that all my doings are wrong and you continously try to put things to my mouth which I have never claimed. And for one reason or another you refuse to prove you claims; you are not discussing here and this is a discussion forum. Apparently you want to stop all discussion about the e factor; I don't know why, your posting history might give somekind of clue.
The data I have used is above, if you know better data, please bring in or point out my errors. Otherwise I continue my search and will ad results to this thread as something new comes in.
gripen
-
And for one reason or another you refuse to prove you claims; you are not discussing here and this is a discussion forum.
I would say he has proved the point on the Wooden Models Gripen.
I would definately say, just as Perkins&Hage states, that Aspect Ratio is a major factor in induced drag formation. I believe the word is "instrumental".
I don't think your qualified to start changing estabilished theories just yet.
Crumpp
-
Originally posted by Crumpp
I would say he has proved the point on the Wooden Models Gripen.
Well, all I see above is that Wood wrote something 1935 and corrections have improved since. And there should be corrections regardless model size.
Originally posted by Crumpp
I would definately say, just as Perkins&Hage states, that Aspect Ratio is a major factor in induced drag formation. I believe the word is "instrumental".
Wow, you have finaly found out that the aspect ratio is a major factor in the induced drag formation:
Cdi = (Cl^2) / (pi * AR * e)
But I think everyone else here allready knew that.
gripen
-
But I think everyone else here allready knew that.
Obviously YOU did not. Your whole argument has been that it is not a factor in determining the efficiency factor. Perkins&Hage say it is along with Wood and the NACA.
Well, all I see above is that Wood wrote something 1935 and corrections have improved since. And there should be corrections regardless model size.
What color is the sky in your world? Can you read?
Crumpp
-
Originally posted by Crumpp
Obviously YOU did not. Your whole argument has been that it is not a factor in determining the efficiency factor. Perkins&Hage say it is along with Wood and the NACA.
Actually I say above that in the case of the tapered wings there seems to be correlation like theory suppose but it is not particularly strong.
But that is not true in every case. If we have no idea about the form of the wing, there is no way to determine e factor from the aspect ratio only.
gripen
-
Originally posted by gripen
Dweeb,
Hm... When I started this thread, only thing I knew was that generalized formulas seem to give quite high value of the e. Since then I have studied quite a lot drag data and so far I can conclude that these generalized formulas really give too high value of the e.
Nope, when you started this thread, you had no idea that the value of the Airplane Efficiency Factor e was affected by Aspect Ratio, infact you wanted others to explain it to you, you even offered a pathetic prize, here is what you actually said:
Originally posted by gripen
In the "Draining E in turns" thread rose a question if the Oswald's efficiency factor (know also as e factor) can be estimated from the unknown wing by knowing only the aspect ratio of that wing. In other words, is it possible to estimate the lift distribution of the wing by knowing only the aspect ratio of it?
So I open a little contest: The first one who can explain if this is possible or not with sensible logic will receive a pdf copy of the RAE report called: "Notes on the Dog Fight" from me. As a bonus prize the winner will also get a flight tested German report on the Bf 109G/AS for performance analysis.
gripen
Then, after you were given sources to refer to, namely: Airplane Performance Stability and Control, by Perkins & Hage you picked up a copy and began posting selective extracts from it pretending you knew the answer the whole time.
Originally posted by gripen
I have not claimed e factor constant, Perkins&Hage makes this quite clear and this can be seen partially from the wind tunnel data as well as flight tested data.
No true Gripen, you actually posted a table of calculations and stated that a constant value of e = 0.8 was better than the formulas posted by the experts… You said:
Originally posted by gripen
It can be seen that generalized formulas result about twice larger average error than the constant value… so it is possible to create a better working general formula than Wood's or NADC.
Gripen, I think I prefer to believe that you are simply going around in circles without realising it because you have such a weak grasp of the subject and that you are trying desperately to comprehend Perkins & Hage fast enough to stay up with this thread, because that only implies you are foolish, the other alternative is that you are simply lying to save face… Your choice?
Originally posted by gripen
I have not claimed that the e factor is same for all WWII fighter planes, again data shows that there is variation but not particularly due to the aspect ratio.
Yes you have, you said that a constant value of 0.8 was better than the equations posted by the experts, and you even backed it up by saying:
Originally posted by gripen
Why the efficiency factor should not be similar among the planes with very similar dimensions?
Which of course is complete baloney, as already explained in previous posts!
Originally posted by gripen
The problem here is that you join in the discussion with the attitude that all my doings are wrong and you continously try to put things to my mouth which I have never claimed.
Nope the real problem is that you have been making absurd claims, and then denying it and lying about it, as your nonsense is shown for what it is.
Originally posted by gripen
And for one reason or another you refuse to prove you claims;
Gripen, you are the only person making claims here, the onus of proof rests entirely with you. You haven’t proven anything here yourself, despite your pompous claims! None of your endless diatribe constitutes proof, it lacks the transparency and rigour of proper academic exposition and actually qualifies as little more than the ranting of a spoilt child trying to gain influence in a gamers board.
Dweeb.
-
Originally posted by Dweeb
Blaah Blaah Blaah
Basicly you are doing here same thing as Crumpp. You post arguments after arguments without proof and you purposedly understand wrong my sayings.
If you want to challenge my calculations, you should post something relevant.
BTW Dweeb's posting history is interesting, quite little posting since 2001. Here (http://www.hitechcreations.com/forums/showthread.php?s=&threadid=66995&perpage=50&pagenumber=2) is a typical example how he pops in and the attitude seem to be allways quite similar as here. Makes me wonder if he has another user name here.
gripen
-
Originally posted by gripen
But that is not true in every case. If we have no idea about the form of the wing, there is no way to determine e factor from the aspect ratio only.
gripen
Gripen, don’t you understand anything you read? Take another look at the equations you have been posting, the aspect ratio is a fundamental part of those equations, you can’t determine e without it, and nobody ever said you could determine e with aspect ratio alone.
Your original post was regarding corrections to e for aspect ratio, and you have already seen in Perkins & Hage that there are corrections for aspect ratio. Glauert has shown that two corrections for aspect ratio influence the drag and the lift curve slope, which has an influence on the efficiency factor first demonstrated by Oswald, and is absolutely based upon the lift and drag coefficients and the aspect ratio… In Perkins & Hage there are graphs of the correction factors as functions of aspect ratio alone, that have been pointed out to you by others. That is enough proof for anyone. You simply couldn’t be more wrong, and you just don’t seem able to see it.
Dweeb
-
Originally posted by gripen
Basicly you are doing here same thing as Crumpp. You post arguments after arguments without proof and you purposedly understand wrong my sayings.
Gripen,
You are the one who makes post after post, all complete nonsense, and you not only refuse to respond to my rational arguments regarding your data, you have not provided any proof for anything you have said, because most of it is contrary to accepted aerodynamic theory. If you want anyone to believe your silly claims, do some original research, have your academic work published and then point us to the publication... The simple fact is, even a first year aero major would laugh at your nonsense, so why don't you just give it up and stop wasting everyone’s time with your complete and utter baloney.
Dweeb
-
Originally posted by Dweeb
You are the one who makes post after...blaah blaah
I'm actually only one who has calculated something here and I have given the sources and so far you have proven nothing wrong on these.
I won't aswer to your arguments because you purposedly understand my sayings wrong and you behave very agressively. For one reason or another you seem keep discussion about determining the e factor as some kind of threat.
The funny thing in your posting history is that it's the most agressive I've seen in this board. I quess the reason for this is that you have another alias for normal use and Dweeb pops up only when you want to be extreme agressive.
gripen
-
I'm actually only one who has calculated something here and I have given the sources and so far you have proven nothing wrong on these.
That is a FLAT OUT LIE.
What do you call this? You know I calculated the "e" factor using every formula you could come up with. Since you did not like the answers you began making up your own theories!! In fact you asked me NOT to be so detailed in my calculations. You couldn't follow it.
Lets see and compare total drag @ 440fps at the Spitfires FTH altitude with the correct data:
BF 274 - 795.641828 @ 440 fps
Total Drag FW-190A5 -718.2845242@ 440 fps
BS354 - No data listed except the A/C was 7 mph slower w/50 bhp more output than the Merlin 66 @ (+18)
BS543 - 810.863789
FW-190A5 - 705.7484202
BS551 - 758.4823884
FW-190A5 - 704.890256
EN524 with 4 bladed prop - 765.4635575
FW-190A5 - 702.4712376
BS310 with 4-bladed prop - 793.0539968
FW-190A5 - 751.8834689
BS310 with 5 bladed prop - 794.0663191
FW-190A5 - 749.7161496
JL165 - done to death
MA 648 Merlin 66 (+18) SU Pump - 838.0549776
FW-190A5 - 804.8037275
BS310 - 792.6552588 Parasitic drag - 691.672177
FW-190A8 - 794.3964059 Parasitic drag - 606.4158622
At 315mph the total drag situation changes completely in the FW-190A8 favor.
JL165 vs FW-190A8 has been done to death, again in the FW-190A8's favor.
MA 648 is the ONLY Spitfire Mk IX to beat the FW-190A8 for drag and here is why:
quote:
--------------------------------------------------------------------------------
External equipment of the above four aircraft was similar with the exception that MA.648 had the new pattern of air intake.
--------------------------------------------------------------------------------
MA648 - 837.1701001 parasitic drag - 751.6212076
FW-190A8 - 869.2121249 parasitic drag - 709.9624067
The parasitic drag is always in the FW-190's favor but the total drag drops to within a few pounds of each other at 315 Mph but never swings in the FW-190's favor.
So ONE example of the Spitfire Mk IX with an experimental air intake beats the FW-190A8 in total drag. The others are behind their FW-190A contemporary the whole way.
That about covers the Spitfire Mk IX's vs. their contemporary FW-190 adversary.
Now this works out the same NO MATTER WHOSE formula you use for "e" factor. At least all the normal accepted ones. It probably won't once Gripen finishes his "theory".
Crumpp
-
So, as a result the Spitfire Mk VIII beats the 190 in total drag then?
Do you know if any of those were clipped?
Oh, edited with the purpose of addition.
You asked about Mach and drag. That is a special category. I have some bit of text about it and can post if you like.
What beats me is that you don't seem to have known of it.
-
Originally posted by gripen
I'm actually only one who has calculated something here and I have given the sources and so far you have proven nothing wrong on these.
Yes but your calculations are all based on specious data, which is why your results are complete nonsense. If you want anyone to take your wind tunnel data seriously, you need to post a lot more information about it, for example, the method of supporting the model and the control of associated turbulent flow problems, and the degree of detail in the model, for example were details such as the radiator details, intakes, pitot heads, radio masts, airspeed and air temperature booms, antenna, guns or gun ports, gun and ammunition door and hatches included on the model? Were any other measures taken to simulate discontinuities in those locations? What corrections were made for Mach number and dynamic pressure calibration? What corrections were made for Reynold’s number? Was the static pressure survey in the tunnel conducted with multiple booms and how were local air disturbances handled? What corrections were made to the Mach number and aerodynamic coefficients for the blockage factor due to the model? What corrections were made to the angle of attack and drag coefficients due to the presence of the tunnel walls? Those are the sort of details you need to be discussing, along with some very self critical estimates of the size of the error bars involved… The burden of proof rests with the person making the claim, and we are still waiting.
Originally posted by gripen
The funny thing in your posting history is that it's the most agressive I've seen in this board. I quess the reason for this is that you have another alias for normal use and Dweeb pops up only when you want to be extreme agressive.
gripen
I assume you mean aggressive… LOL! I can’t believe I’m arguing with someone who can’t even spell his accusations correctly! Gripen, why don’t you stop pouting and wipe your teary eyes and try to provide some of the proof you have been asked for, instead of just repeating the same old nonsense over and over again, and finally resorting in desperation, to conspiracy theories?
Originally posted by gripen
I won't aswer to your arguments because you purposedly understand my sayings wrong and you behave very agressively. For one reason or another you seem keep discussion about determining the e factor as some kind of threat.
Gripen, your spelling is almost as bad as your aerodynamics. Three incorrect words in a single sentence, I wonder if you have even finished high school?
However, the real reason you won’t answer my arguments is because you can’t! I have challenged you to provide the proper transparency and rigour required to transform your arguments from the nonsense they are. The burden of proof always and without exception, rests with the person making the claim. It is therefore your responsibility to demonstrate that your data and the resulting calculations are valid, which is impossible to do because “Performance estimates based on model test results are unsatisfactory because of the large and uncertain corrections necessary to make the results correspond to the free flight of a full sized airplane.”
The only things you have proven successfully in this thread, is that your data and thus the resulting calculations are meaningless, that you haven’t owned your copy of Perkins & Hage for long enough to understand it properly yet, and that your general level of education is somewhere between the 7th and 8th grade... Congratulations!
Dweeb
-
Dweeb, don't be a dweeb and start picking on gripen about spelling details if you have nothing moreworthy to promote.
BTW, it is most common for non-english Europeans to spell some words as "address" and "aggressive" with a single d or g, since these are international terms, such as in Swedish, German and such.
Now, if you're an American, just imagine the Brits go picking on you. LOL, Torpedo, Tomato, etc...:D
Now, you can fight about wind tunnels as you like. There is however a reason for the application of wind tunnels, bear that in mind. After all, they are the ultimate simulation before a flight test, - highly adjustable, and modifications are quick in for a test.
The end-effect is what matters. If you know the thrust and the weight, the rest is up to the aerodynamics, and real life this time!
So, if a Spifire is faster than a 190 with the same thrust at a given alt there must be less total drag, - and vice versa.
But in a wind tunnel there may be even less error in the thrust department, note that !
We have data showing most in the 190's favour maybe (?). However some Spits seem to claw themselves in front.
So why?
-
Originally posted by Angus
Dweeb, don't be a dweeb and start picking on gripen about spelling details if you have nothing moreworthy to promote. BTW, it is most common for non-english Europeans to spell some words as "address" and "aggressive" with a single d or g, since these are international terms, such as in Swedish, German and such. Now, if you're an American, just imagine the Brits go picking on you. LOL, Torpedo, Tomato, etc...:D
Point taken.
Originally posted by Angus
Now, you can fight about wind tunnels as you like. There is however a reason for the application of wind tunnels, bear that in mind. After all, they are the ultimate simulation before a flight test, - highly adjustable, and modifications are quick in for a test.
Any designer will tell you there is only one reason they use models in wind tunnels… cost! Airplane design by building, testing, changing and retesting models is essentially a confession of ignorance of the designer, it is an attempt to make design by trial and error, less expensive than the millions involved in working with full size aircraft. No designer would ever dream of claiming that their model test results are likely to be close in terms of performance, that’s not what they are intended for. They save a lot of money, and a lot of big mistakes in terms of balance, dynamic stability, and integration, in the very early stages of the design process. More and more money has been spent refining the techniques and corrections to the point where the wind tunnel testing of models has become an essential part of aircraft design that can save billions on modern projects. In the 1940s model testing was in its infancy, and designers couldn’t use those tests for performance prediction with any hope of worthwhile results and they still don’t. That’s why development programs include the full range of tests, from scale models initially, to full size model testing, to prototype tests all before the aircraft ever gets near a runway, and eventually every new aircraft is subject to a huge range of flight tests, and even then, performance estimates will be changing as the tests evolve. I don’t normally like to be so blunt, but using model tests the way Gripen has, is just plain silly, and he wouldn’t believe it even if Kurt Tank told him in person ;-)
Originally posted by Angus
But in a wind tunnel there may be even less error in the thrust department, note that !
We have data showing most in the 190's favour maybe (?). However some Spits seem to claw themselves in front.
So why?
Sorry Angus, I don’t understand what you are asking? Scale model tests are never conducted with engines installed, so thrust isn’t a factor.
Dweeb
-
We have data showing most in the 190's favour maybe (?). However some Spits seem to claw themselves in front.
There are a lot of factors. Temperature, humidity, barometric pressure, power variance, finish of the aircraft, etc...
There is a natural variance in all manufactured products. Could be the riveter was having a good day and just did an exceptional job on that particular plane. Or maybe the finish guy just got a really poor paint job with lots of orange peel etc. Anyway you get the idea.
So all planes will vary within a couple of percentage points from a center mark. You get a really good one paired with a really bad one and your results can be completely different from the truth. That is why I did ALL the Spitfire Mk IX's I could find good verifiable data on.
That sound right to you, Dweeb? You probably have some good insight into this phenomenon.
I think if we had a wider base of FW-190 data we could get an even better picture. I have quite a few more flight test graphs. Think I might run the data through them and see what comes of it.
Crumpp
-
After some search I found the source for Wood's formula:
Karl D. Wood (1935):"Technical Aerodynamics"
The formula:
ew = 1/(0,97+0,033*R)
is for rectangular wings only.It is directly derived from Glauert corrections and it is not supposed to work with other wing shapes (tapered etc.) nor it won't give directly the efficiency factor of the whole airframe. Wood describes estimation method for a whole airfame but it is based also on drag coefficient and fuselage shape. Also NADC formula seems to be for a straight untapered wing only.
Basicly these formulas are not really generalized formulas but for certain wing shapes and these can't be used to estimate e factor of the whole airframe except in the case of the rectangular flying wing.
As for additional info here is some e factor values from VL archives:
VL Pyörremyrsky 0,74 (wind tunnel data, 1/6 model)
VL Puuska 0,77 (apparently calculation)
VL Humu 0,80 (calculation)
gripen
-
is for rectangular wings only.It is directly derived from Glauert corrections and it is not supposed to work with other wing shapes (tapered etc.) nor it won't give directly the efficiency factor of the whole airframe. Wood describes estimation method for a whole airfame but it is based also on drag coefficient and fuselage shape. Also NADC formula seems to be for a straight untapered wing only.
Do you even READ your Perkins&Hage??
Crumpp
-
Crumpp,
What about Perkins&Hage?
gripen
-
Originally posted by Badboy
However, the formulae presented by Professor Wood is for a taper ratio of 0.57 which yields an almost elliptical lift distribution, which is why it fits the experimental data so closely over a wide range of examples. Most WWII aircraft had a taper ratio close to that value (Me109 was approx 0.52) because the designers knew about the benefits of elliptical lift distribution and that it could be achieved quite closely with a wing of that taper, that’s why a formulae that only includes aspect ratio could still be of such good practical use.
Here is the scan from the "Technical Aerodynamics" (p.120):
(http://personal.inet.fi/koti/soon.moro/woodf.jpg)
The formula is not for tapered wings with certain taper ratio but simply for rectangular wings and here is a quote from the page 36 (above the part Badboy quoted in another thread):
"In the practical airplane calculations the graphs for tapered wings may be assumed to lie between the graphs for rectangular and elliptical wings. For ratios of tip chord to root chord between 0.2 and 0.6, the tapered wing has practically the same characteristics as the elliptical wing."
Shortly the formula is not supposed to be a generalized formula to calculate e factor for wings as Badboy presents above, it just helps to estime e for the wing with the above figure. For entire airframe the effect of the fuselage and other surfaces should be added.
gripen
-
A tapered wing never equals the elliptical wing in lift distribution, however, it simulates it to within acceptable limits, giving a tremendous bonus in manufacture may I add.
A tapered wing may also present a much better aspect ratio.
However, the ellipse always pays the least induced drag penalty for the given total lift. Even if a similar tapered wing has a more favourable aspect ratio!
That's why it's partially called an aerobatic-wing
Then, there is the chord.
I am curious of the benefit of wide chord.
Would that be lift while banking?
-
Gripen,
I’ve only just realized where we appear to have our lines crossed and what has caused some of the confusion that has crept into this thread, and I think I can clear it up. Throughout this thread you have been questioning why the formulae posted seem to give values of the e factor that are too high, and that don’t correspond to empirical data… You said this in a previous message…
Originally posted by gripen
All I say above is that generalized formulas by Wood and NADC seem to give too high value of the e factor for WWII fighters and this conclusion is backed up by empirical data.
I didn’t realize what was happening until I saw your recent posts, and I think I can explain what you are seeing. We have been referring to Oswald’s efficiency factor, (Airplane Efficiency factor) and we have been using the character e to represent it, which is what normally happens, but there is a catch... There are two values associated with it and they are not the same, and they don’t measure the same things. Take another look at the explanation I posted previously:
Originally posted by Badboy
The number normally represented by the character e in induced drag calculations was originally known as Oswald’s efficiency factor, and his original paper is available for download from the NACA report server. More commonly it has a component of parasite drag lumped in with it and is just called the airplane efficiency factor and can be estimated depending on the aspect ratio, taper ratio, sweep angle and twist. Theoretically an elliptical wing would have an efficiency factor of 1, meaning that it will have a coefficient of induced drag close to the theoretical maximum. Even though it is a function of aspect ratio, sweepback angle, taper ratio, camber, Mach number and twist, the largest influence on the wing of a WWII fighter with very little sweep or twist comes from aspect ratio and taper ratio and so there are approximate formulae for estimating e that only include aspect and taper ratio, and even more approximate methods that only include aspect ratio. The important thing you must appreciate is that they are only approximate. But better than just assuming a constant value for every aircraft..
You might remember that, but the important point is where I explain that the airplane efficiency factor has a component of parasite drag lumped in with it. In your copy of Wood’s book you can see this, and I’ll point out the references to Wood and Perkins & Hage in a moment. Firstly, let me explain what happens… If you plot a graph of parasite drag and induced drag you can see that they both vary with the coefficient of lift. This is expressed mathematically in 2-82 at the top of page 93 of Perkins & Hage. Normally the two terms that vary with lift are lumped together, and when you do that, you are combining the induced drag from the wing, and induced drag due to a component of the parasite drag resulting from the lifting capability of the fuselage and the tail, in fact from the aircraft as a whole. In the normal parabolic drag polar, the accounting is done that way so that all the lift related components can be lumped together in a single term like the one in 2-83 on the same page. The problem is that they are not always combined in the same way, and when the wing is being treated in isolation, during airfoil theory for example, you may see the induced drag for the wing expressed using a value for e that does not include the effects of the lift dependant component of the parasite drag for the fuselage or tail, or anything other than the induced drag of the wing. Wood does both in his book and he uses the term ew with the subscript w to indicate a value for the wing only, and he uses e by itself for the airplane efficiency factor for the aircraft as a whole. If you look at empirical data, such as the drag polar from a flight test, you will see the later value, if you see a drag polar for an airfoil you will see the former value. I believe this has caused much of the confusion in this thread, because we have been discussing two different things, and of course the formulae posted were for the wing, which is why you have noticed that those values are too high. I hope that clears up that misunderstanding.
Now I would like to address your following point:
Originally posted by gripen
The formula is not for tapered wings with certain taper ratio but simply for rectangular wings and here is a quote from the page 36 (above the part Badboy quoted in another thread):
"In the practical airplane calculations the graphs for tapered wings may be assumed to lie between the graphs for rectangular and elliptical wings. For ratios of tip chord to root chord between 0.2 and 0.6, the tapered wing has practically the same characteristics as the elliptical wing."
What Wood is saying in your quote is that the value for all wings will fall between the two graphs shown below, with elliptical wings at one extreme and rectangular at the other.
(http://www.badz.pwp.blueyonder.co.uk/images/wood.jpg)
You can see from this graph, that the value for wings with various other taper ratios can be determined simply by choosing a line between them. So in effect you can have a generalized formula for any taper. But a formulae that only yields a value for the wing alone. However, Wood goes on to say in the part of the quote that you appear to have left off:
" Glauert gives corrections for tapered wings as a function of the amount of taper, but this refinement is believed to be not justified in practice."
When this thread started, you were only asking for someone to explain how e could be estimated from aspect ratio, however, if you wish to include taper ratio as well, you can do it using a slightly different formulae, the formulae is:
e = (2083T – 1083)/(33R + 970) - 2.083T + 2.083
Where R is the aspect ratio and T is the taper ratio between 1 for a rectangular wing and 0.52 equivalent to an elliptical wing (you can only use values between 1 and 0.52 in this equation). So for example, if you calculate for an aspect ratio of 6 and a taper ratio of 0.8 you get a value for e = 0.92. For an aircraft with a particular fuselage shape and size that could end up being closer to 0.83 for the whole aircraft.
Glauert also published a very nice graph that allows you to determine e from both the aspect ratio and from the taper ratio, I can post it if that would be helpful?
But it is better to work from drag polars derived from flight test data.
Hope that helps…
Badboy
-
That was the point, without the math, I made a long time ago..
-B
Originally posted by Badboy
Gripen,
I’ve only just realized where we appear to have our lines crossed and what has caused some of the confusion that has crept into this thread, and I think I can clear it up. Throughout this thread you have been questioning why the formulae posted seem to give values of the e factor that are too high, and that don’t correspond to empirical data… You said this in a previous message…
...
Badboy
-
Originally posted by Badboy
...however, if you wish to include taper ratio as well, you can do it using a slightly different formulae, the formulae is:
e = (2083T – 1083)/(33R + 970) - 2.083T + 2.083
This gives a better estimate than a formula you originally used which was just for rectangular wing. Anyway, most WWII fighters had some amount of washout which should be counted if we want to determine e from the wing geometry and Glauert gives induced drag corrections for washout too.
Otherwise the effect of the fuselage seems to around 10% if calculated from the drag polars like in your example.
gripen
-
Gripen,
You keep just glossing over:
" Glauert gives corrections for tapered wings as a function of the amount of taper, but this refinement is believed to be not justified in practice."
And
"For studies requiring a higher degree of accuracy, lifting-surface theories have been used [as opposed to lifting line theory], but generally it has been found that the additional complexity of these methods has not sufficiently improved the predictions to warrant common use." (NACA Report 921 - Theoretical Symmetric Span Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form).
People have told you from the begining of this thread. Before it was discovered you were using a wooden model and passing it off as the real thing.
The new formula changes the numbers but not the results. Do the math and you will see.
Crumpp
-
My dear Gentlemen.
A new year has started.
I looked over many a post, and I see some mistakes I've done.
I've seen some, others have done.
(luckily, much more, hehe)
I've seen arguments that always go in circles.
So, sometimes it's a sour apple to swallow, but how about putting some things aside and proceed.
I am proud of being "in the club" with so many of you fellows, for you promote things to learn from. And there is no end to learning.
So,
For instance, Crumpp and Gripen.
Sometimes you are fighting about almost nothing. And it always goes a bit bitter, instead of perhaps, useful, if you know what I mean.
Can't say I don't know this, just look at me and Izzy :D
Well, anyway, may these threads live on.....
-
And a PS, since you are into the effect of tapered wings vs totally square ones, I might be able to help.
(Will have to go book-diving though)
The (LIFT-) INDUCED drag will, if my memory serves me, be some 10% less with a moderately tapered wing.
The only 2D wingform with less induced drag is the ellipse.
-
" Glauert gives corrections for tapered wings as a function of the amount of taper, but this refinement is believed to be not justified in practice."
Well, actually the induced drag part of the Wood's formula for rectangular wings is based on Glauert corrections. In addition Anderson (http://naca.larc.nasa.gov/reports/1937/naca-report-572/) (1936) used Glauert corrections for his study and pointed out:
"The calculated values of the pitching-moment coefficient at zero lift , the aerodynamic center position, the angle of the zero lift , and the lift-curve slope are generally in good agreement with the tests values (table VI). The agreement of the pitching moment coefficient at zero lift and the aerodynamic-center position, which are calculated from the basic and additional lift distributions, respectively, indicate that the theoretical lift distributions must also agree reasonably well with the actual distributions "
This is also confimed by Abbott and von Doenhoff (1949) "Theory of Wing Sections" and actually even Badboy above offers to post some kind "very nice chart" on Glauert corrections (thanks but I've got a copy of Glauert's book allready). Glauert corrections are still fundamental and well accepted part of the wing theory.
Shortly Wood's book is many ways out dated, as an example he wrotes about wind tunnel testing with scaled down models:
".. but in general the smaller the tunnel the larger the corrections for scale and turbulence which must be made to have results applicable to full-size airplanes.Such corrections are uncertain in the present state of art and introduce large errors into the results."
Just about ten years later Pope (1947, "Wind tunnel testing") pointed out that lift curve slope can be well estimated with the models regardless the size of the model becauce the slope of the lift curve remains about same regardless the Reynold's number. In addition Pope's book contain a nice listing of wind tunnels around the world including Germany; there was no large wind tunnel in Germany during WWII. BTW Pope's books are still available, I think latest was published in nineties.
Regarding the subject of this thread, it has been pointed out that the Wood's formula for rectangular wings as well the NADC formula can't be directly used as generalized formulas and there is no way to estimate e factor of the wing or entire airframe accurately from the aspect ratio only.
Originally posted by Crumpp
The new formula changes the numbers but not the results. Do the math and you will see.
There is no washout included in the formula so it won't give accurate results for the planes with washout. Besides, there should be a viscous part of drag included for elliptical wing (or wing with optimal taper ratio, 0,52 or what ever). This means that the e factor won't be 1 even in the optimal case due to viscous part of the drag rise (see Wood figure 32).
gripen
-
Just about ten years later Pope (1947, "Wind tunnel testing") pointed out that lift curve slope can be well estimated with the models regardless the size of the model becauce the slope of the lift curve remains about same regardless the Reynold's number. In addition Pope's book contain a nice listing of wind tunnels around the world including Germany; there was no large wind tunnel in Germany during WWII. BTW Pope's books are still available, I think latest was published in nineties.
And we are back to square one.
Your saying Badboy is full of crap? Wooden models CAN be used for accurately determining the e factor when compared with actual aircraft?
Not for drag estimation, Gripen. Call HTC and ask Hitech.
BTW the Drag polars for the FW-190 where done at a large wind tunnel in FRANCE. As has been said many times in this thread. Since I have the entire report and it includes pictures....
It's hard to claim it was a wooden model Gripen.
Regarding the subject of this thread, it has been pointed out that the Wood's formula for rectangular wings as well the NADC formula can't be directly used as generalized formulas and there is no way to estimate e factor of the wing or entire airframe accurately from the aspect ratio only.
What reality do you live in? Have you read anything Badboy or anyone else has posted?
" Glauert gives corrections for tapered wings as a function of the amount of taper, but this refinement is believed to be not justified in practice."
This has got to be an English as a second language mistake. Surely you are smarter than that or you are just trolling. Please break out a dictionary and look up the meaning of the words in this statement. It is obvious you do not understand them.
Crumpp
-
Originally posted by Crumpp
Your saying Badboy is full of crap? Wooden models CAN be used for accurately determining the e factor when compared with actual aircraft?
I have not said nothing about content of the Badboy despite he has posted to this thread with double identities (Dweep being extremely aggressive) and the formula he claimed to generic was actually not. Anyway, unlike you, Badboy is capable to discuss about the subject rational way and able to admit his errors. Basicly you have posted over 100 times to this thread without a single relevant post so far.
What I say above is that the Wood's opinions about the corrections were outdated soon after his book came out. Actually one major breaktrough happened just in same year (NACA 547 (http://naca.larc.nasa.gov/reports/1937/naca-report-547/)) that's actually one of the Pope's sources.
Originally posted by Crumpp
Not for drag estimation, Gripen. Call HTC and ask Hitech.
We are not talking about the total drag here but the drag rise due to change of the lift coefficient (induced and viscous). And this can be estimated from the models with good accuracy given the corrections are right. A wooden model might be cleaner than a real plane but the lift characters as well as viscous drag change due to increasing AoA are about same regardless the size of the model.
Originally posted by Crumpp
BTW the Drag polars for the FW-190 where done at a large wind tunnel in FRANCE.
I've got a report as well as pictures too. The point was that the Fw data in the drag chart posted above for Ta 152 is certainly based on model testing and probably the values for the Fw 190 too because the chart contains detailed data for parts of the airframe as well as claims for interference drag which is a clear indication of model testing.
Originally posted by Crumpp
Have you read anything Badboy or anyone else has posted?
I have read his postings as well as checked his as well as your sources.
Originally posted by Crumpp
" Glauert gives corrections for tapered wings as a function of the amount of taper, but this refinement is believed to be not justified in practice."
This has got to be an English as a second language mistake. Surely you are smarter than that or you are just trolling. Please break out a dictionary and look up the meaning of the words in this statement. It is obvious you do not understand them.
Just read Anderson as well as Perkins&Hage as well as Abbott&vonDoenhoff... They all use Glauert corrections. Even Wood used Glauert correction for rectangular wing, Wood's book is simply outdated and actually he uses words "believed to be not justified". Besides even Badboy claims above "Glauert also published a very nice graph".
You are the one who is trolling here.
gripen
-
I've got a report as well as pictures too. The point was that the Fw data in the drag chart posted above for Ta 152 is certainly based on model testing and probably the values for the Fw 190 too because the chart contains detailed data for parts of the airframe as well as claims for interference drag which is a clear indication of model testing.
Not the report I am talking about. I am refering to the actual polar plots of the FW-190.
We are not talking about the total drag here but the drag rise due to change of the lift coefficient (induced and viscous). And this can be estimated from the models with good accuracy given the corrections are right. A wooden model might be cleaner than a real plane but the lift characters as well as viscous drag change due to increasing AoA are about same regardless the size of the model.
We are talking about Drag ESTIMATION, Gripen. Again give HTC a call. Badboy has already told you on this one as well. Specifically the efficiency factor as part of the induced drag formula.
You need to read the reports you post.
Wood's opinions about the corrections were outdated
It's not just Wood, Gripen. Even the report you posted about the influence of the shape of the windtunnel says the same thing!
"For studies requiring a higher degree of accuracy, lifting-surface theories have been used [as opposed to lifting line theory], but generally it has been found that the additional complexity of these methods has not sufficiently improved the predictions to warrant common use." (NACA Report 921 - Theoretical Symmetric Span Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form).
Again do the math.
Besides even Badboy claims above "Glauert also published a very nice graph".
Ok. Barney the Purple Dinosaur publish es some nice coloring books too.
Your taking this completely out of context. Badboy's chart showed that the range was so narrow that correction made little difference. Just like he explained in his post.
Twhich is a clear indication of model testing.
BS. You would not know clear indications of model testing if it came up and bit you. You've shown yourself to be about as knowledgeable on aeronautics as anyone else on this board inspite of the fact you attempt to pass yourself off as an expert.
Now your attempting to justify your wooden model antics with the 1/6th Spitfire polars.
Crumpp
-
Originally posted by Crumpp
Not the report I am talking about. I am refering to the actual polar plots of the FW-190.
There is not much need to post further data for this thread, it has been already pointed out that there is no way to estimate e factor accurately from the aspect ratio only. But if you have finally something relevant for this thread, please post.
Originally posted by Crumpp
We are talking about Drag ESTIMATION, Gripen. Again give HTC a call.
To be exact, we are talking about the estimation of e factor of the whole airframe ie the change of the drag (induced and viscous) when the lift coefficient changes, not about total drag. Pyro and Hitech are most wellcome to join in discussion anytime if you need them to tell you what Pope and others have allready pointed out.
Originally posted by Crumpp
Badboy has already told you on this one as well.
Well, the problem here is that Badboy has continously misundestood what Wood and others actually wrote. The formula he originally posted was not so called gereric formula and another example is Glauert; Glauert never made corrections to the e factor because he published his work couple years before Oswald and Glauert corrections are for Prandtl's lifting line theory ie for induced drag only.
Originally posted by Crumpp
It's not just Wood, Gripen. Even the report you posted about the influence of the shape of the windtunnel says the same thing!
Let's see the whole quote, keep in mind that mentioned modifications for lifting line theory actually mean Glauert corrections:
"Lifting-line theory in the past has been so modified and extended so that the characteristics of wings having no sweep, moderate to high aspect ratio, and any taper ratio can be determined readily with good accuracy. For studies requiring a higher degree of accuracy, lifting-surface theories have been used [as opposed to lifting line theory], but generally it has been found that the additional complexity of these methods has not sufficiently improved the predictions to warrant common use."
So actually the quote says that Glauert corrections for the lifting line theory work well. Besides I have not used Glauert or Wood, my analyses are strictly based on drag polars which even according to Badboy is the right method.
Originally posted by Crumpp
Badboy's chart showed that the range was so narrow that correction made little difference.
Well, if we ad the effect of the washout, increased viscous drag for optimal taper ratio, and the fuselage, the values for the P-51 and Fw 190 are very close to the values taken from the drag polars ie somewhere around 0,75-0,8. Funny thing is that even Raymer uses constant value of e 0,75 in his simplified analysis.
Originally posted by Crumpp
BS. You would not know clear indications of model testing if it came up and bit you.
There was no large wind tunnels in the Germany so it's safe to say that Fw data is at least partially based on small models.
gripen
-
There is not much need to post further data for this thread, it has been already pointed out that there is no way to estimate e factor accurately from the aspect ratio only.
Only in your mind. People who know what they are talking about use them all the time. Why don't ask an aeronautical engineer?
To be exact, we are talking about the estimation of e factor of the whole airframe ie the change of the drag (induced and viscous) when the lift coefficient changes, not about total drag.
You did not understand what Badboy pointed out did you? There are TWO e factors's. Only Oswalds includes an element of viscous drag.
my analyses are strictly based on drag polars which even according to Badboy is the right method.
What a blatent lie.
1. Your Spitfire is a drag polar of a wooden model which you are trying to pass off as representative of the actual plane.
2. You do not have a drag polar for the FW-190. In fact most of the calculations you have done have been off half baked theories of your own. Such as the ones you did off Lednicer's lift distribution chart.
There was no large wind tunnels in the Germany so it's safe to say that Fw data is at least partially based on small models.
Guess you totally missed the pictures that have already been posted of full sized aircraft hanging in Rechlin wind tunnels.
There is no point in continuing this conversation. You have nothing and are simply trying to "get" the last word in and try and save a little face. You know most people will only read the last post in a long thread and are desperately hoping to salvage something and recreate the fantasy of Gripen the expert.
Crumpp
-
Originally posted by Crumpp
People who know what they are talking about use them all the time.
I quess this claim is in the same leaque as the "wet lifting area".
Originally posted by Crumpp
You did not understand what Badboy pointed out did you? There are TWO e factors's. Only Oswalds includes an element of viscous drag.
There is just one e factor and it was introduced by Oswald 1931. It's for entire airframe and includes induced and viscous drag. There is no other e factor. Besides I have told you nearly two months ago that lift distribution analysis (like Glauert corrections) do not count viscous part of the drag rise (see my posts 11-12-2004 09:17 PM and 11-15-2004 02:42 PM) so I have known perfectly well that e factor contains induced and viscous drag long before Badboy posted something on this.
Originally posted by Crumpp
1. Your Spitfire is a drag polar of a wooden model which you are trying to pass off as representative of the actual plane.
2. You do not have a drag polar for the FW-190.
The models can be used perfectly well to determine e factor as pointed out by Pope and others. The Fw data sheet gives the values of the K and that is all I need for drag polar.
Originally posted by Crumpp
Guess you totally missed the pictures that have already been posted of full sized aircraft hanging in Rechlin wind tunnels.
The tunnel picture which you posted was the 8m tunnel A3 in the LFA (Brunswick) ie not a large scale tunnel.
Originally posted by Crumpp
There is no point in continuing this conversation.
So far you have posted nothing relevant to this thread, so this might be your best idea so far in this thread.
gripen
-
Gripen says:
The models can be used perfectly well to determine e factor as pointed out by Pope and others.
Pope says NOTHING about comparing wooden models to actual aircraft of a different type Gripen. Pope says the CONCLUSIONS of Wooden model testing GENERALLY give good agreement with the SAME aircraft. In other words the general conclusions hold true but not the exact numbers!!
You just can't seem to understand that.
Badboy has pointed it out.
Badboy says"
Any designer will tell you there is only one reason they use models in wind tunnels… cost! Airplane design by building, testing, changing and retesting models is essentially a confession of ignorance of the designer, it is an attempt to make design by trial and error, less expensive than the millions involved in working with full size aircraft. No designer would ever dream of claiming that their model test results are likely to be close in terms of performance, that’s not what they are intended for. They save a lot of money, and a lot of big mistakes in terms of balance, dynamic stability, and integration, in the very early stages of the design process. More and more money has been spent refining the techniques and corrections to the point where the wind tunnel testing of models has become an essential part of aircraft design that can save billions on modern projects. In the 1940s model testing was in its infancy, and designers couldn’t use those tests for performance prediction with any hope of worthwhile results and they still don’t. That’s why development programs include the full range of tests, from scale models initially, to full size model testing, to prototype tests all before the aircraft ever gets near a runway, and eventually every new aircraft is subject to a huge range of flight tests, and even then, performance estimates will be changing as the tests evolve. I don’t normally like to be so blunt, but using model tests the way Gripen has, is just plain silly, and he wouldn’t believe it even if Kurt Tank told him in person ;-)
The tunnel picture which you posted was the 8m tunnel A3 in the LFA (Brunswick) ie not a large scale tunnel.
Looks to me like they have an actual plane hanging in it.
(http://www.onpoi.net/ah/pics/users/503_1104873746_109windtunnel.jpg)
So far you have posted nothing relevant to this thread
Your right you have not posted anything relevant.
Crumpp
-
Guys,
Please leave me out of your discussions.
Many thanks.
Badboy
But on a passing note Gripen, one last source for you, that includes some software I've been using to estimate e as a function of taper, aspect ratio, mid-chord sweep and Mach number. The software uses spanwise load distributions calculated by linearised lifting-surface theory. If you want to include every aspect of the geometry and fusilage you may find computational fluid dynamics of more value.
http://www.esdu.com/graphics/dataitem/74035a.htm
Here is a sample output from the software I just ran as an example, based on the F6F. You can see that in this case the taper ratio is only just over 0.52 and the last column in the output data can be inverted for e = 0.997:
**************************************************
ESDU International plc
Program A7435
ESDUpac Number: A7435V10
ESDUpac Title: Subsonic lift-dependent drag due to the trailing
vortex wake for wings without camber or twist.
Data Item Number: 74035
Data Item Title: Subsonic lift-dependent drag due to the trailing
vortex wake for wings without camber or twist.
ESDUpac Version: 1.0 Issued April 1996.
(See Data Item for full input/output specification and interpretation.)
**************************************************
INPUT DATA
==========
F6F
A=5.51,MID-CHORD SWEEP=5 deg.,TAPER RATIO=0.5275
INPUT FILE test.txt
INPUT DATA ERRORS
-----------------
No error detected
OUTPUT DATA
===========
----------------------------------------------------------------------
M A n Ln (deg) Taper BetaA AtanLh 1+Delta
----------------------------------------------------------------------
.100 5.510 .500 5.00 .527 5.48 .48 1.003
.150 5.510 .500 5.00 .527 5.45 .48 1.003
.200 5.510 .500 5.00 .527 5.40 .48 1.003
.250 5.510 .500 5.00 .527 5.34 .48 1.003
.300 5.510 .500 5.00 .527 5.26 .48 1.003
.350 5.510 .500 5.00 .527 5.16 .48 1.003
.400 5.510 .500 5.00 .527 5.05 .48 1.003
.450 5.510 .500 5.00 .527 4.92 .48 1.003
.500 5.510 .500 5.00 .527 4.77 .48 1.002
------------------------ END OF OUTPUT ---------------------------
Hope that helps...
Badboy
-
Originally posted by Crumpp
Pope says NOTHING about comparing wooden models to actual aircraft of a different type Gripen. Pope says the CONCLUSIONS of Wooden model testing GENERALLY give good agreement with the SAME aircraft. In other words the general conclusions hold true but not the exact numbers!!
Actually he says that the size of the model does not affect lift curve slope despite the reynolds number changes. It does not matter how the results are used.
Originally posted by Crumpp
Badboy has pointed it out.
Actually that was "Dweeb" who said for example that Glauert made corrections for Oswald's efficiency and other quite questionable claims...
Basicly the model testing has allways been most common way to test properties of the airframe before flight testing and the corrections have been good enough for accurate analysis since forties.
Originally posted by Crumpp
Looks to me like they have an actual plane hanging in it.
As pointed out earlier, the wing tips are out of the tunnel area and lift coefficient analysis are practically impossible. The most advenced wind tunnel in the LFA was the high speed wind tunnel A2 which was a 2,8m tunnel and that was the tunnel were the germans tested advanced planes (with scale models).
Originally posted by Crumpp
Your right you have not posted anything relevant.
Hm...I have posted several drag polars, sources, a true flight test, calculation methods etc. Besides original question of this thread has been solved long time ago, no need to post more data.
Badboy,
As you can read from my latest reply to your post, I'm not looking trouble with you.
Regarding computer analysis I can say that atleast the CFD analysis underestimates viscous drag. You might allready know from where this comes...
gripen
-
Actually that was "Dweeb" who said for example that Glauert made corrections for Oswald's efficiency and other quite questionable claims...
:rofl :rofl
Basicly the model testing has allways been most common way to test properties of the airframe before flight testing and the corrections have been good enough for accurate analysis since forties.
Great! WE can inform the aerospace industry that they can save millions on test flights!
Gripen says the data from their 1/6th scale models is exact. Therefore we can just skip the test flight stage and go right to production!
:aok
As pointed out earlier, the wing tips are out of the tunnel area and lift coefficient analysis are practically impossible. The most advenced wind tunnel in the LFA was the high speed wind tunnel A2 which was a 2,8m tunnel and that was the tunnel were the germans tested advanced planes (with scale models).
Yeah ok. Get some glasses.
Again though. The Focke Wulf report uses the windtunnel in :
(http://www.onpoi.net/ah/pics/users/503_1104877639_polartitle.jpg)
Which is definately large enough.
Hm...I have posted several drag polars, sources, a true flight test, calculation methods etc. Besides original question of this thread has been solved long time ago, no need to post more data.
BS You have posted your "theories" on calculating e from lift distribution graphs, used the drag polar from a wooden model (Spitfire), and ONE drag polar from the an unknown source on the 109. Don't sell yourself to dearly there Gripen.
Crumpp
-
Originally posted by Badboy
Here is a sample output from the software I just ran as an example, based on the F6F. You can see that in this case the taper ratio is only just over 0.52 and the last column in the output data can be inverted for e = 0.997:
The Glauert correction gives about same result for taper ratio 0,52 so it seems that this software calculates just induced drag.
Originally posted by Crumpp
Great! WE can inform the aerospace industry that they can save millions on test flights!
They allready know this and they have allready saved millions with model testing.
Originally posted by Crumpp
The Focke Wulf report uses the windtunnel in :
Chalais-Meudon was a large scale tunnel (16x8m) but it was also a low speed tunnel and the germans lost it summer 1944. All high speed tunnel testing of Fw was certainly done with models and probably all testing of the Ta 152 series.
gripen
-
with models and probably all testing of the Ta 152 series.
That's a really great piece of fiction. Please PROVE IT. Based on the actual report, your full of it.
In the world according to Gripen, the Germans could never do any testing! You always try and come up with something to support your agenda.
They allready know this and they have allready saved millions with model testing.
And now according to your findings they can just eliminate the whole test flight stage of the development process!!
:lol
Crumpp
-
Originally posted by gripen
The Glauert correction gives about same result for taper ratio 0,52 so it seems that this software calculates just induced drag.
Yes, it just determine the subsonic lift-dependent drag due to the trailing vortex wake for planar straight-tapered wings in inviscid subsonic flow. The theory from which the data was derived is linearised and relates to inviscid flow. These restrictions are known to be relatively unimportant for most practical purposes in calculating the lift and lift-dependent drag due to the trailing vortex wake provided the flow remains fully attached over the wing surface. The paper also says that in applying this software it should be borne in mind that camber and twist and the presence of a body, engine nacelles etc, all affect trailing vortex drag to a greater or lesser degree. Of those mentioned the body effect is considered to be the most important and is stated to be generally detrimental owing to the distortion of the spanwise loading in its vicinity. This is the effect that Wood is accounting for in applying his fuselage factor.
Hope that helps...
Badboy
-
Thanks Badboy.
Crumpp
-
Originally posted by Crumpp
That's a really great piece of fiction. Please PROVE IT. Based on the actual report, your full of it.
As noted earlier it's up to you to prove that there was a large scale tunnel in the Germany.
Originally posted by Crumpp
In the world according to Gripen, the Germans could never do any testing!
Great logic again. The Germans had several advanced smaller scale high speed tunnels, which they used for testing.
Originally posted by Crumpp
And now according to your findings they can just eliminate the whole test flight stage of the development process!!
I'm not saying that but the e factor (change in the induced and viscous drag due to change of the lift coefficient) can be estimated with good accuracy regardless the size of the model.
Originally posted by Badboy
Yes, it just determine the subsonic lift-dependent drag due to the trailing vortex wake for planar straight-tapered wings in inviscid subsonic flow.
Yep, so the viscous part of the drag is missing and we can't calculate e factor for the wing.
Originally posted by Badboy
Of those mentioned the body effect is considered to be the most important and is stated to be generally detrimental owing to the distortion of the spanwise loading in its vicinity. This is the effect that Wood is accounting for in applying his fuselage factor.
Wood's system to estimate the effect of the fuselage is based on NACA 236 (http://naca.larc.nasa.gov/reports/1927/naca-report-236/) which is a compilation of tests results on various fuselage, nacelle etc. data from year 1927. In most of these tests there was no tail surfaces in the fuselage and the effect of the interference drag is not counted.
gripen
-
As noted earlier it's up to you to prove that there was a large scale tunnel in the Germany.
How many times do I have to tell your dense self. FRANCE Gripen, the report says FRANCE! In fact it names the wind tunnel. You do not need a high speed wind tunnel to do drag estimations or polars. OBVIOUSLY!!!
I'm not saying that but the e factor (change in the induced and viscous drag due to change of the lift coefficient) can be estimated with good accuracy regardless the size of the model.
No Gripen it cannot. The CONCLUSIONS in a generalized fashion can be determined from a wooden model NOT the exact numbers!!
Crumpp
-
Originally posted by Crumpp
How many times do I have to tell your dense self. FRANCE Gripen, the report says FRANCE!
As noted several times the germans lost France summer 1944, therefore it is safe to say that at least Fw data for the Ta 152 is based on tests with scale models and even the data for the Fw 190s seem to be at least partially based on model testing because it it contains detailed drag values for different parts of the airframe and claims about interference drag.
Besides the Chalais-Meudon was a low speed tunnel (IIRC about 45m/s) and Fw data sheet values for K are for much higher speed, about 80m/s at climb and well over 100m/s at high speed flying. So probably all values in the Fw datasheet are based on model testing.
Originally posted by Crumpp
You do not need a high speed wind tunnel to do drag estimations or polars.
No one has stated otherwise here. But the the most advanced German planes under developement reached transonic (compressibility) speeds so the need for high speed tunnels was obivious for drag and other analysis. Therefore they developed smaller and advanced high speed tunnels instead large low speed tunnels.
Originally posted by Crumpp
No Gripen it cannot. The CONCLUSIONS in a generalized fashion can be determined from a wooden model NOT the exact numbers!!
Pope is very clear on this, lift curve as well as drag rise due to lift coefficient change can be determined with good accuracy despite changes in the value of the Reynolds number. Besides we don't need to be very accurate here; wind tunnel data as well as flight tests indicate that the value of the e factor for the WWII fighters was typically somewhere around 0,75 just like Raymer uses in his simplified analysis.
gripen
-
As noted several times the germans lost France summer 1944, therefore it is safe to say that at least Fw data for the Ta 152 is based on tests with scale models and even the data for the Fw 190s seem to be at least partially based on model testing because it it contains detailed drag values for different parts of the airframe and claims about interference drag.
You need to study up more on the Ta-152 program. Prototypes were flying in 1943 Gripen as were the FW-190A9 prototypes.
Pope is very clear on this, lift curve as well as drag rise due to lift coefficient change can be determined with good accuracy despite changes in the value of the Reynolds number. Besides we don't need to be very accurate here; wind tunnel data as well as flight tests indicate that the value of the e factor for the WWII fighters was typically somewhere around 0,75 just like Raymer uses in his simplified analysis.
AS accurate as your trying to be in your comparison of a 1/6th wooden model spitfire to the actual FW-190 aircraft??
You cannot use a wooden model in comparison to an actual aircraft for exact numbers!
Post your reference for Pope!
I guarantee he does not say what you are trying to have us believe.
Crumpp
-
Originally posted by Crumpp
You need to study up more on the Ta-152 program. Prototypes were flying in 1943 Gripen as were the FW-190A9 prototypes.
There were no completed Ta 152 prototypes before summer 1944. Besides it's up to you to prove full scale wind tunnel testing of the Ta 152.
Originally posted by Crumpp
AS accurate as your trying to be in your comparison of a 1/6th wooden model spitfire to the actual FW-190 aircraft??
...
You cannot use a wooden model in comparison to an actual aircraft for exact numbers!
As noted several times earlier, only you are making comparisons here. This thread is about determination of the e factor not about comparisons.
Originally posted by Crumpp
Post your reference for Pope!
I guarantee he does not say what you are trying to have us believe.
A. Pope (1947): "Wind-Tunnel Testing" p. 274-275:
"It has been observed that (Fig. 7:8) that, when Cl^2 for a given airplane is plotted against total drag coefficient Cdt, the graph is nearly a straight line. Further, since we may write.. [formula]...it becomes apparent that the slope of the line dCd/dCl^2 may be used to find e. (See also Sect. 5:2) Fortunately the slope of this line is practically independent of Reynolds number and a wind tunnel test may hence be used to determine full scale e."
(http://personal.inet.fi/koti/soon.moro/pope.jpg)
Basicly the size of the model has very little effect to the slope of the line. Therefore scaled models can be well used to determine e. It can be also seen that the differences in the Cd0 do not cause differences in the slope of the line.
gripen
-
Since you have been busted before taking things out of context and only including the portions that support YOUR claim but not including the conclusions which are contrary, Gripen, please just post the page/paragraph.
There were no completed Ta 152 prototypes before summer 1944. Besides it's up to you to prove full scale wind tunnel testing of the Ta 152.
http://users.belgacom.net/aircraft1/avion1/64.html
* Tank continued to tweak the inline-powered designs, resulting in the "Ta-152" series, with work along this line begun in late 1942. The "Ta" stood for "Tank", in honor of his contributions to the Reich. A confusing number of different Ta-152 variants were considered or built in prototype form in 1943 and 1944, converging on two types, the short-wing "Ta-152C" and the long-wing "Ta-152H", where the "H" stood for "Hoehenjaeger (High Altitude Fighter)."
http://www.faqs.org/docs/air/avfw190.html#m5
Common misconception Gripen. The Ta-152 prototypes as well as the FW-190A9 prototypes were flying in 1943.
Crumpp
-
Originally posted by Crumpp
Since you have been busted before taking things out of context and only including the portions that support YOUR claim but not including the conclusions which are contrary, Gripen, please just post the page/paragraph.
The source is claimed, get the book if you don't want to believe the direct quote and the graph.
It's up to you to prove that Pope is wrong.
Originally posted by Crumpp
Common misconception Gripen. The Ta-152 prototypes as well as the FW-190A9 prototypes were flying in 1943.
From Baugher's site (http://www.csd.uwo.ca/~pettypi/elevon/baugher_other/ta152.html):
"The first Ta 152H prototypes were completed in the summer of 1944."
Common misconception is in your sources; while the Fw 190C prototypes were used in the developement of the Ta 152 and were flying 1943 in addition to Fw 190 V32 (Ta 153 test plane), these all had standard Fw 190 wing at that time. First true Ta 152 (W.Nr. 150001, CW+CA) was completed June 1944.
Besides your source contains clear errors, there was no prototypes of the Ta 152A and B.
gripen
edit: There seem to had been be couple Ta 152B protypes completed 1945 but no Ta 152A prototypes.
-
Common misconception is in your sources; while the Fw 190C prototypes were used in the developement of the Ta 152 and were flying 1943 in addition to Fw 190 V32 (Ta 153 test plane), these all had standard Fw 190 wing at that time. First true Ta 152 (W.Nr. 150001, CW+CA) was completed June 1944.
My source is:
http://www.schifferbooks.com/newschiffer/book_template.php?isbn=0764318764
April of 1944. Plenty of time to use any windtunnel in Europe.
besides your source contains clear errors, there was no prototypes of the Ta 152A and B
Another wrong statement.
Try FW-190V20.
t's up to you to prove that Pope is wrong.
Popes not wrong. You are. I am certain you are either quoting out of context or as you did before just plain leaving the part you don't want out. If you were not doing so it would not be a big deal to just post the page and underline it. I will get the book, post it myself.
Crumpp
-
Originally posted by gripen
Yep, so the viscous part of the drag is missing and we can't calculate e factor for the wing.
They state that even though the theory from which the data was derived is linearised and relates to inviscid flow, those restrictions are known to be relatively unimportant for most practical purposes in calculating the lift and lift-dependent drag. It does give you the e value for the wing, within those limitations.
Originally posted by gripen
Wood's system to estimate the effect of the fuselage is based on NACA 236 which is a compilation of tests results on various fuselage, nacelle etc. data from year 1927. In most of these tests there was no tail surfaces in the fuselage and the effect of the interference drag is not counted.
Yep, as Wood states on page 66 of his book. However, all I meant was, that they both used the same approach in terms of finding a value for the wing separately from the fuselage, not that they used exactly the same methods to do so, sorry, I should have made that clearer.
If you would like a copy of the report and the software, let me know?
Badboy
-
Originally posted by Crumpp
quote: besides your source contains clear errors, there was no prototypes of the Ta 152A and B
Another wrong statement.
Try FW-190V20.
Not just the V20 (W.Nr. 0042, TI+IG) but also the Fw190 V19 (W.Nr 0041) and Fw190 V21 (W.Nr 0043, TI+IH).
These 3 a/c were produced at Adelheide for the testing of the planned Ta 152A series.
However, the 2 planned prototypes for the Ta 152 A-1 (TA 152 V1, W.Nr 250 001 and Ta152 V2, W.Nr 250 002) with the new 19.5 m2 wing were cancelled.
Harmann's Ta 152 book p 23
-
Originally posted by Crumpp
April of 1944. Plenty of time to use any windtunnel in Europe.
April 1944 is not 1943. Besides it's up to you to prove that the Ta 152 prototype was tested in the full scale wind tunnel, not that some kind of prototype existed june or april 1944.
Originally posted by Crumpp
Popes not wrong. You are. I am certain you are either quoting out of context or as you did before just plain leaving the part you don't want out. If you were not doing so it would not be a big deal to just post the page and underline it. I will get the book, post it myself.
Well, you are most wellcome to do so if you don't want to believe me.
Originally posted by Badboy
They state that even though the theory from which the data was derived is linearised and relates to inviscid flow, those restrictions are known to be relatively unimportant for most practical purposes in calculating the lift and lift-dependent drag. It does give you the e value for the wing, within those limitations.
As noted earlier in this thread, also the viscous drag of the wing increase simultaneously with the rise of the induced (lift related) drag when the Cl (or AoA) increases. And to determine e factor we need to count this part of drag rise too.
As an example we can use Wood's formula for rectangular wing and compare it to Glauert corrected data for same wing type (assuming AR 6):
Wood => e=0,86 (induced and viscous)
Glauert => u=0,94 (only induced)
Note that I use Anderson's induced drag factor u which is used same way as e.
gripen
-
April 1944 is not 1943. Besides it's up to you to prove that the Ta 152 prototype was tested in the full scale wind tunnel, not that some kind of prototype existed june or april 1944.
No it is not Gripen. As I said before I have the full report. April is plenty of time to fly for an hour and spend a day or two at the wind tunnels in France.
Crumpp
-
Originally posted by Crumpp
No it is not Gripen. As I said before I have the full report.
Well, then just post the evidence that the Ta 152 was tested in the large scale wind tunnel.
gripen
-
Gripen says:
Well, then just post the evidence that the Ta 152 was tested in the large scale wind tunnel.
Gripen says:
The source is claimed, get the book if you don't want to believe the direct quote and the graph.
Crumpp says:
If you want the report go to the archives and dig through the index card list and find your own copy.
My credibility is intact, Gripen. I am not the one who has been busted quoting selected parts of data to advance an agenda. Or advancing ridiculus theories of my own.
Crumpp
-
Originally posted by Crumpp
If you want the report go to the archives and dig through the index card list and find your own copy.
So I quess this claim is again in the same category as the "wet lifting area" ie you are talking about a thing which does not exist.
Originally posted by Crumpp
My credibility is intact, Gripen. I am not the one who has been busted quoting selected parts of data to advance an agenda. Or advancing ridiculus theories of my own.
Unlike you I have identified my sources and anyone can check them if needed. Besides from above anyone can find a good example of your quoting in the case of the NACA 921.
gripen
-
Unlike you I have identified my sources and anyone can check them if needed. Besides from above anyone can find a good example of your quoting in the case of the NACA 921.
Well do it.
Only source I have not released is this report.
Gripen says:
I quess this claim is again in the same category as the "wet lifting area" ie you are talking about a thing which does not exist.
Crumpp says:
Exactly. Why is that? Because the taper ratio for the Spitfire is impossible to nail down since the wet lifting area in no way matches the elliptical shape of the wing due to the wing twist destroying the benefits of the elliptical tips. That leaves a wide margin that delivers correct values that are meaningless in the real world. Very open to data manipulation.
No Gripen I was talking about the lift distribution and confused the terminology. Unlike you I have never represented myself as an expert.
Lets Check out your claims:
1. All aircraft have the same e factor ".8"!
2. You did not realize that the differences in values comes from an element of parasitic drag. Niether did I until Badboy pointed it out.
3. Your ridiculous calculations off Lednicers lift distribution chart. The one three aeronautical engineers working in the field could not figure out how you could make any conclusions about efficiency factor.
4. Your repeated insistence in using skewed data even when undisputable proof is laid before you.
5. Your insistence then that the aircraft finish was not representative even when presented with undisputable proof.
6. When the math did not go your way you started this ridiculous thread offering a "prize" looking for help for the other thread!
Posting my documentation will NOT prove anything to you. You will continue with your agenda no matter what anyone else says or produces. Badboy has tried to square you away and you argue with him still!
Only thing it would do is give you the benefit of my time and money spent researching it.
Crumpp
-
Regardless of the conclusions they draw, Gripen and Badboy are capable of well-mannered and intelligent discourse and reasoning out their differences.
Crumpp and Dweeb could use a serious lesson from Miss Manners. You come across as frothing zealots who continually twist others words and meanings so that even if you have a point it's hard to take you seriously.
I know nobody gives a rats bellybutton what I think but I just wanted to let you all know how you come across to the random passer-by.
Stick that in your pipe!
-
Originally posted by Crumpp
Well do it.
Let's look it again, below is Crumpp's version (his post 01-04-2005 12:50 AM) of the NACA 921 which was supposed to support Wood's claim on validity of the Glauert corrections for tapered wings:
"For studies requiring a higher degree of accuracy, lifting-surface theories have been used [as opposed to lifting line theory], but generally it has been found that the additional complexity of these methods has not sufficiently improved the predictions to warrant common use." (NACA Report 921 - Theoretical Symmetric Span Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form)"
Let's see again the whole quote, keep in mind that mentioned modifications for lifting line theory actually mean Glauert corrections:
"Lifting-line theory in the past has been so modified and extended so that the characteristics of wings having no sweep, moderate to high aspect ratio, and any taper ratio can be determined readily with good accuracy. For studies requiring a higher degree of accuracy, lifting-surface theories have been used [as opposed to lifting line theory], but generally it has been found that the additional complexity of these methods has not sufficiently improved the predictions to warrant common use."
So actually the quote says that Glauert corrections for the lifting line theory work well which is exactly opposite Crumpp wanted to prove.
Originally posted by Crumpp
Only source I have not released is this report.
Well, I don't see any proof here that such report on the Ta 152 exist.
Originally posted by Crumpp
No Gripen I was talking about the lift distribution and confused the terminology. Unlike you I have never represented myself as an expert.
Well, I wonder what are you doing in this thread then because you seem to have no clue what you are talking about.
Originally posted by Crumpp
Lets Check out your claims:
1. All aircraft have the same e factor ".8"!
Such claim I have not done.
Originally posted by Crumpp
2. You did not realize that the differences in values comes from an element of parasitic drag. Niether did I until Badboy pointed it out.
As pointed out above I have told you nearly two months ago that lift distribution analysis (like Glauert corrections) do not count viscous part of the drag rise (see my posts 11-12-2004 09:17 PM and 11-15-2004 02:42 PM) so I have known perfectly well that e factor contains induced and viscous drag long before Badboy posted something on this.
Originally posted by Crumpp
3. Your ridiculous calculations off Lednicers lift distribution chart. The one three aeronautical engineers working in the field could not figure out how you could make any conclusions about efficiency factor.
As pointed out earlier, there is certainly a strong correlation between lift distribution an efficiency factor and that was what I calculated with good results. Or do you have some kind of evidence that there is no correlation between e factor and lift distribution?
Originally posted by Crumpp
4. Your repeated insistence in using skewed data even when undisputable proof is laid before you.
I wonder what are you talking about here.
Originally posted by Crumpp
5. Your insistence then that the aircraft finish was not representative even when presented with undisputable proof.
Nothing to do with this thread and that case I was just quoting a source you self use.
Originally posted by Crumpp
6. When the math did not go your way you started this ridiculous thread offering a "prize" looking for help for the other thread!
It has been pointed out in this thread that there is no way to calculate e factor from the aspect ratio only.
Originally posted by Crumpp
Posting my documentation will NOT prove anything to you.
Only thing it would do is give you the benefit of my time and money spent researching it.
So far you have not posted anything relevant for this thread so I wonder why you keep posting.
gripen
-
Oh, come on.
I mean, is this a duel or what?
-
et's look it again, below is Crumpp's version (his post 01-04-2005 12:50 AM) of the NACA 921 which was supposed to support Wood's claim on validity of the Glauert corrections for tapered wings:
You just told a story Gripen. You posted nothing I quoted!
Typical.
Crump says:
Your ridiculous calculations off Lednicers lift distribution chart. The one three aeronautical engineers working in the field could not figure out how you could make any conclusions about efficiency factor.
Gripen says:
I had no wind tunnel data to calculate e factor of the Fw 190. But Lednicer calculated span loadings and here is estimated e factors calculated by taking 8 samples (semispan fractions 0,2-0,9) and measuring distance to the elliptical span loading:
Crump says:
2. You did not realize that the differences in values comes from an element of parasitic drag. Niether did I until Badboy pointed it out.
Gripen says:
If we look these numbers and compare them to generalized formulas (like Wood's or the one in the Zigrat's sheet), we can see that except the case of the Spitfire, generalized formulas seem to give somewhat higher values (around 10%, and in the case of the Spitfire the accuracy seem to be more or less accidental). We can also see that this very limited data set supports assumption that the e factor decreases when the aspect ratio increases. In addition we can also see that taper ratios are somewhat lower than assumed in the Wood's formula and except the case of the Bf 109, the planes have some washout (I don't know if the generalized formulas assumed washout).
In spite of his being told:
Joeblogs says:
Getting back to Grippen's point, though, Oswald's e is used to correct for departures from eliptical planform and is therefore not just a function of the aspect ratio.
Even I said on page 2 of this thread:
A correction is made for the variation of parasitic resistance with angle of attack and nonelliptical wingloading by including in the induced drag term a factor e, called the "airplane efficiency factor". The correction is thus assumed to be proportional to CL^2.
He continues on and on about the differences formulas calculated values.
Another pearl of wisdom from Gripen.
Gripen says:
Overall the exact speed (if below Mach 0,5) is not important if we know the Cl and/or value of the K.
Crumpp says:
If you want to compare planes it certainly is important. Nobody cares that when the FW-190 and the Spitfire have the same CL the spitfire's wing is more efficient. What matters is under the SAME conditions of flight.
And yet another:
Crumpp says:
Lets Check out your claims:
1. All aircraft have the same e factor ".8"!
Gripen says:
The problem is that that the WWII fighters generally had lower e factor which means that the both formulas give too high estimates and therefore also a constant value works better in most cases.
I could be here all night cutting and pasting!
Lastly for g00b:
Crumpp and Dweeb could use a serious lesson from Miss Manners.
Dweeb is Badboy's pseudonym.
Crumpp
-
Originally posted by Crumpp
You just told a story Gripen. You posted nothing I quoted!
Anyone can check this from Crumpp's posting (http://www.hitechcreations.com/forums/showthread.php?s=&postid=1517128#post1517128).
Originally posted by Crumpp
Even I said on page 2 of this thread:
Anyone can check (http://www.hitechcreations.com/forums/showthread.php?s=&postid=1388941#post1388941) that there Crumpp just quotes Oswald's report (http://naca.larc.nasa.gov/reports/1933/naca-report-408/index.cgi?page0001.gif) which I linked earlier.
Otherwise Crumpp just continues his cut and paste game. It has been pointed out that so called generalized formulas were actually formulas for rectangular wing only and can't be used for entire airframe nor other wing types
gripen
-
Crumpp just quotes Oswald's report which I linked earlier.
Sure you linked it Gripen but you sure did not understand it!
Crumpp
-
Originally posted by Crumpp
Sure you linked it Gripen but you sure did not understand it!
So why I have pointed out several times during this thread that also viscous drag rise should be counted for e factor?
Besides just two days ago you announced (http://www.hitechcreations.com/forums/showthread.php?s=&postid=1517626#post1517626) that "There are TWO e factors's".
Basicly you continously quote sources without understanding them, Oswald and NACA 921 are good examples.
gripen
-
Crumpp says:
You did not understand what Badboy pointed out did you? There are TWO e factors's. Only Oswalds includes an element of viscous drag.
Yep, Gripen. And there are two e factors.
Badboy says:
I didn’t realize what was happening until I saw your recent posts, and I think I can explain what you are seeing. We have been referring to Oswald’s efficiency factor, (Airplane Efficiency factor) and we have been using the character e to represent it, which is what normally happens, but there is a catch... There are two values associated with it and they are not the same, and they don’t measure the same things. Take another look at the explanation I posted previously:
ONE
Badboy says:
you may see the induced drag for the wing expressed using a value for e that does not include the effects of the lift dependant component of the parasite drag for the fuselage or tail, or anything other than the induced drag of the wing.
TWO
Badboy says:
The number normally represented by the character e in induced drag calculations was originally known as Oswald’s efficiency factor, and his original paper is available for download from the NACA report server. More commonly it has a component of parasite drag lumped in with it and is just called the airplane efficiency factor and can be estimated depending on the aspect ratio, taper ratio, sweep angle and twist. Theoretically an elliptical wing would have an efficiency factor of 1, meaning that it will have a coefficient of induced drag close to the theoretical maximum. Even though it is a function of aspect ratio, sweepback angle, taper ratio, camber, Mach number and twist, the largest influence on the wing of a WWII fighter with very little sweep or twist comes from aspect ratio and taper ratio and so there are approximate formulae for estimating e that only include aspect and taper ratio, and even more approximate methods that only include aspect ratio. The important thing you must appreciate is that they are only approximate. But better than just assuming a constant value for every aircraft..
Thanks for proving my point again!
Crump says:
Sure you linked it Gripen but you sure did not understand it!
Crumpp
-
Originally posted by Crumpp
Yep, Gripen. And there are two e factors.
As noted earlier there is just one e factor which is for entire airplane and it includes induced drag rise and viscous drag rise just like noted in the Oswald's study. In the Wood's book, the author separated this to two pieces, ew for wing and ef for fuselage.
The factor which include just induced drag is not the e factor, in the Anderson's study it's called the factor u.
Regarding Badboy's postings, he has allready admited missunderstandings. However, unlike you Badboy is capable to discuss about this rational way as others have noted. And I don't want to mix him in this discussion like you continously do.
In addition I noted a funny thing above when you wrote about exact speeds in the Fw drag chart. If the shape of the lift/drag curve below compressibility speeds (say mach 0,5) is not practically constant, it means that measurements in the low speed tunnel like Chalais-Meudon (around 45m/s) are worthless.
gripen
-
Regarding Badboy's postings, he has allready admited missunderstandings.
Umm Looks to me like your the one with the misunderstandings, Gripen.
When are you going to admit them?
In addition I noted a funny thing above when you wrote about exact speeds in the Fw drag chart. If the shape of the lift/drag curve below compressibility speeds (say mach 0,5) is not practically constant, it means that measurements in the low speed tunnel like Chalais-Meudon (around 45m/s) are worthless.
You should let Kurt Tank know!!
Crumpp
-
Just to summarise the technical aspects of this discussion…
Anyone reading this thread could be forgiven for wondering why the value of e should be so important. Why argue about the difference between something as small as 0.8 and 0.85 for example, when it only has a small influence on the overall drag coefficient. But before we get into this, let’s just apply a crude reality check to see if we have a realistic range for the fighters we are interested in? This diagram:
(http://www.badz.pwp.blueyonder.co.uk/images/Oswald2.jpg)
taken from NACA 408 (http://naca.larc.nasa.gov/reports/1933/naca-report-408/) shows that Oswald’s estimate agrees with the values that arise from approximate equations solely based on aspect ratio, for example, values between 0.85 and 1 for a cantilever monoplane. It is worth noting that he also quotes values between 0.95 and 1 just for a wing on its own, which is similar to approximate values produced earlier in this thread for a wing also.
Well, let’s put that in terms of air combat, and look at the difference that would make to an aircraft at the very bottom of Oswald’s range 0.85, and one even lower, say 0.8 corresponding to a value at the high end of the range of average values Gripen posted from drag polars for various WWII fighters earlier in this thread.
Well, here is a diagram that shows the difference that these two values would have on the sustained turn rate of the same aircraft. Firstly, it would make very little difference at all to any other performance characteristics, the top speed for example being only 0.4mph different (and hardly distinguishable on the chart). You can see from the diagram that there is only 0.6 degrees per second difference (less than 3%) in the sustained turn rate, and no difference in the sustained turn radius, or any of the instantaneous values.
(http://www.badz.pwp.blueyonder.co.uk/images/EMe3.jpg)
Alternatively, the pilot in the aircraft with an e = 0.8 could choose to match the turn rate of the less draggy counterpart, but to do so he would have to lose altitude in the turn at the rate of 260ft/min.
That’s a bout the size of it, not a decisive advantage by any means and because in a real engagement, that difference is small enough to be overwhelmed by other factors, such as pilot ability, fuel or other ordnance loads, or the significant differences between the dissimilar aircraft more likely to have been involved in real combat.
Lastly, the diagram also includes the equations used in a modern context to describe Oswald’s efficiency factor, and they demonstrate three things. Firstly, that the definition of e as Oswald intended, includes the factors that have been discussed here in that context correctly by Gripen, but separately on occasions by me :) Those factors are the value for the wing alone, indicated by the Greek character delta in those equations, and the value for the fuselage indicated by the character k. Some modern analytical methods estimate those values together or separately, while practical methods provide the value of e directly from the slope of a graph produced from either wind tunnel or flight test data.
I think that just about sums it up…
Hope that helps put things in perspective…
Badboy
-
Good Post.
Definately puts things in persepective.
Crumpp
-
Originally posted by Badboy
Anyone reading this thread could be forgiven for wondering why the value of e should be so important. Why argue about the difference between something as small as 0.8 and 0.85 for example, when it only has a small influence on the overall drag coefficient.
Well, the point of this thread was clear up the methods to determine the e for performance calculations and we (me and Badboy) have a consensus now here.
Originally posted by Badboy
But before we get into this, let’s just apply a crude reality check to see if we have a realistic range for the fighters we are interested in? This diagram:
(http://www.badz.pwp.blueyonder.co.uk/images/Oswald2.jpg)
taken from NACA 236 (http://naca.larc.nasa.gov/reports/1927/naca-report-236/) shows that Oswald’s estimate agrees with the values that arise from approximate equations solely based on aspect ratio, for example, values between 0.85 and 1 for a cantilever monoplane. It is worth noting that he also quotes values between 0.95 and 1 just for a wing on its own, which is similar to approximate values produced earlier in this thread for a wing also.
Just a little correction. The chart linked above is from NACA 408 (http://naca.larc.nasa.gov/reports/1933/naca-report-408/) which is Oswald's report where he introduced the e factor, mentioned NACA 236 is the reference Wood used for his partial e factor for fuselage.
Generally the values in the Oswald's chart are probably a bit on high side; the value of e must be always be less than 1 as Wood notes. Practical maximum is probably around 0,95 for the flying wings. Here is some values (http://web.usna.navy.mil/~dfr/flttst1b.pdf) for modern planes from Navy web site:
"The value e can be estimated for a particular model of aircraft by various methods.
One of the simplest is by comparison with aircraft of similar configuration. For most
aircraft e has a value of about 0.6 to 0.8. Some values of e for various aircraft as
measured at Mississippi State University are
Bellancer Crusair (low-wing) 0.55
Learstar (twin engine mid-wing) 0.67
Cessna 170 (high-wing) 0.74
RJ-5 Glider (high-wing) 0.79"
Originally posted by Badboy
Lastly, the diagram also includes the equations used in a modern context to describe Oswald’s efficiency factor, and they demonstrate three things. Firstly, that the definition of e as Oswald intended, includes the factors that have been discussed here in that context correctly by Gripen, but separately on occasions by me :) Those factors are the value for the wing alone, indicated by the Greek character delta in those equations, and the value for the fuselage indicated by the character k. Some modern analytical methods estimate those values together or separately, while practical methods provide the value of e directly from the slope of a graph produced from either wind tunnel or flight test data.
A practical comparison can be done with the Spiteful (no washout) wind tunnel data and calculation for entire airframe using the estimated value of K 0,01 (viscous drag) from Perkins&Hage and Glauert correction factor 0,011 for taper ratio about 0,5 (induced drag):
Wind tunnel => e=0,81
Calculated => e=0,84
Which is a pretty good result given the generalized form of the estimation. The effect of the wash out is much more difficult to calculate using the Glauert corrections because the effect varies with the AoA. Maybe the software used by Badboy could used for that case?
Originally posted by Crumpp
You should let Kurt Tank know!!
Tank certainly knew that below compressibility speeds the shape of the drag polar is practically constant, therefore the polars measured below these speeds (say mach 0,5) are comparable. Basicly if you are saying something else, it means that the polars measured at low speed tunnel are not good for other speeds than the speed used in the measurement (about 45m/s in the case of the Chalais-Meudon).
gripen
edit: The link to the Navy site corrected.
-
Tank certainly knew that below compressibility speeds the shape of the drag polar is practically constant, therefore the polars measured below these speeds (say mach 0,5) are comparable. Basicly if you are saying something else, it means that the polars measured at low speed tunnel are not good for other speeds than the speed used in the measurement (about 45m/s in the case of the Chalais-Meudon).
I am saying your calculated value is not close to the value you get off the polar, Gripen.
Crumpp
-
Originally posted by Crumpp
I am saying your calculated value is not close to the value you get off the polar, Gripen.
Hm... You said that there is something wrong in my statement regarding the speed in which the value of the K for climb in the Fw chart was given assuming that there is about linear stage in the Cd/Cl^2 curve:
"Overall the exact speed (if below Mach 0,5) is not important if we know the Cl and/or value of the K".
If this statement is not true, the polar which is measured in the low speed tunnel can't be used for other speeds.
If this statement is true, the polar which is measured in the low speed tunnel can be used for any practical flying speed below compressibility speeds.
gripen
-
Originally posted by gripen
Just a little correction. The chart linked above is from NACA 408 which is Oswald's report where he introduced the e factor
Thanks for picking that one up, that's what I get for not reading it back before posting... I've corrected it.
Badboy
-
A practical comparison can be done with the Spiteful (no washout) wind tunnel data and calculation for entire airframe using the estimated value of K 0,01 (viscous drag) from Perkins&Hage and Glauert correction factor 0,011 for taper ratio about 0,5 (induced drag):
Wind tunnel => e=0,81
Calculated => e=0,84
Here is a good example. Your saying the Spiteful has a more efficient wing than the FW-190A not be just a little but a rather large margin. Your calculations using "K" from Focke Wulf's data say the e factor is around .78!!
Lets look at the Spiteful's wing:
http://www.fortunecity.com/marina/manatee/272/spiteful.html
Now we all know wingtip efficiency can be very easily manipulated thru structural design. Designers were very aware of the benefits of wing efficiency in the late 1930's and were familiar with how to manipulate the design to attain an efficient wing.
The FW-190 and the Spiteful have almost the exact same wingtip. As you know wingtips are extremely important to wing efficiency.
Both the FW-190 and the Spiteful have sharp cornered wingtips. Figure (c) in this diagram. As you can see the wingtip design is not far behind an elliptical tip.
http://www.centennialofflight.gov/essay/Theories_of_Flight/Reducing_Induced_Drag/TH16G6.htm
Another tool designers use manipulate efficiency is twist.
They include (1) planform taper to obtain an elliptic planform, used for the Spitfire wing, which was remarkably elliptic; (2) a geometric twist and/or aerodynamic twist to obtain elliptic lift distribution; or (3) a combination of all of these methods.
The FW-190 wing was twisted 2 degrees specifically to achieve an elliptical distribution. As Lednicer points out it was not far behind the Spitfire and neither was the P51. The lift distribution chart puts the wingtips almost equal. All three aircraft are very close with the Spitfire having a slight advantage.
Now as we all know the Spitfires elliptical distribution was purposely destroyed by it designers to lessen the harsh stall characteristics inherent with a perfect ellipse.
The other problem with elliptical wings is the stall characteristics. It is much safer to design an airplane so that the wing stalls first at the root, leaving the outer portion of the wing, (where the ailerons are) still flying. An elliptical wing however, will tend to stall uniformly all along the span (see the diagram below.) The "fix" for this situation is washout, but that will reduce the theoretical gains in induced drag. Therefore, we are unlikely to see a great resurgence in the use of elliptical wings, except in situations where appearance dictates.
http://142.26.194.131/aerodynamics1/Drag/Page8.html
So I think we can all agree that .88 is about right for the Spitfire. This is backed up by EVERYONE's calculations.
Another factor is aspect ratio. The FW-190 had the advantage over the Spiteful with a higher aspect ratio.
6.02 = FW-190
5.81 = Spiteful
The efficiency factor e and wing span are physical factors that may be controlled by proper design. A plane with a longer span wing (higher aspect ratio) has less induced drag and, therefore, greater efficiency.
So frankly your calculations do not make sense putting the FW-190 so far behind the Spitfire.
Spiteful - Same wingtip, no twist, lower aspect ratio, e factor = .84
FW-190 - Same wingtip, 2 degree twist, higher aspect ratio, .78!!
Aeronautical engineer using the same data calculates the FW-190's e factor to be .87.
Every formula I have used puts its the FW-190's in the same ballpark. .86-.87.
Crumpp says:
I am saying your calculated value is not close to the value you get off the polar, Gripen.
Is a true statement.
Now you might be tempted to state that e factor includes an element of parasitic drag. Well we both know the FW-190 has less parasitic drag than the Spitfire so I find that hard to imagine that the viscous drag portion of "e" would make a .10 difference!
As for the polars themselves, Gripen.
The polar include multiple pages and multiple aircraft configurations.
The include prop mounted, no prop, low power high power, gear up, gear down, flaps up, flaps down, cooling gills open cooling gills closed etc...
Even graphs for different hardpoint racks. Your numbers do not correspond to any of them.
Each configuration and speed has it's own graph.
I would look to your math before I looked to FW documents. Your numbers do not make sense.
Crumpp
-
(http://deephousepage.com/smilies/deadhorse.gif)
-
Originally posted by Crumpp
Your saying the Spiteful has a more efficient wing than the FW-190A not be just a little but a rather large margin. Your calculations using "K" from Focke Wulf's data say the e factor is around .78!!
This has absolute nothing to do with my calculations; the value of the K (1,24) at climb speed for the Fw 190 results value of the e 0,78. There is no other answer for this value of the K.
The Spiteful wind tunnel data gives a very well behaving Cd/Cl^2 curve, practically linear:
(http://personal.inet.fi/koti/soon.moro/Spiteful.jpg)
Again there is no other answer for the value of the e because the line is so linear despite what ever calculation method is used.
Originally posted by Crumpp
The FW-190 wing was twisted 2 degrees specifically to achieve an elliptical distribution. As Lednicer points out it was not far behind the Spitfire and neither was the P51.
There was no washout in the Spiteful and taper ratio is near optimal, while Lednicer's analysis prove that the Fw 190 had too much washout for optimal lift distribution:
(http://personal.inet.fi/koti/soon.moro/sp.jpg)
Originally posted by Crumpp
Another factor is aspect ratio. The FW-190 had the advantage over the Spiteful with a higher aspect ratio.
6.02 = FW-190
5.81 = Spiteful
The e factor of the tapered wing decreases slightly when the aspect ratio increases, you can test this yourself with formulas given above by Badboy and you can also find them from Perkins&Hage. Again this has nothing to do with my calculations.
gripen
-
There was no washout in the Spiteful
Exactly. The whole point Gripen is:
1. The OPTIMAL shape is elliptical. If the physical construction of the wing is not elliptical then you must manipulate the wing structure to achieve an elliptical distribution. This is done by twist. The Spiteful does not have elliptical wing construction. It has NO twist either to move it closer to elliptical.
The FW-190 does have a 2 degree twist to move it closer to elliptical. Just like the P51. The P51 and the FW-190 were given twist to bring them closer to elliptical. The Spitfire was given 2 degrees of twist to move it away from perfect elliptical due to the stall.
taper ratio is near optimal,
Yeah, OK.
Optimal taper ratio is elliptical construction, Gripen.
http://www.centennialofflight.gov/essay/Theories_of_Flight/Reducing_Induced_Drag/TH16G6.htm
Since the lift distribution is nearly elliptical, the chord distribution should be nearly elliptical for uniform Cl's. Reduced lift or t/c outboard would permit lower taper ratios.
http://www.desktopaero.com/appliedaero/wingdesign/wingparams.html
Planform taper is the reduction of the chord length and thickness as one proceeds from the root (near the fuselage) to the tip section (at the wing tip) so that the airfoil sections also remain geometrically similar. (A planform is the shape of the wing as one looks down on it from above.)
http://www.centennialofflight.gov/essay/Theories_of_Flight/Reducing_Induced_Drag/TH16.htm
I find it hard to believe. Especially since the inclusion of taper ratio is considered
not to make a large difference at all and is generally considered unnecessary for most applications.
In fact the largest contributor to e factor is wingtip construction.
The wing-tip shape, being at the point where the tip vortices are produced, appears to be of more importance in minimizing tip vortex formation and thus minimizing induced drag. Taper and twist are perhaps of greater importance in dealing with the problem of stalling.
http://www.centennialofflight.gov/essay/Theories_of_Flight/Reducing_Induced_Drag/TH16.htm
A taper ratio of 0.4 most closely approximates the elliptical.
The FW-190's Taper Ratio is according to you - .44
The most optimal would be .40.
Since both Aspect Ratio and Taper very very slightly factor into e factor I would have to say the Aspect Ratio advantage of .21 for the FW-190 trumps the .04 Taper Ratio advantage of the Spiteful.
So I have to call a big BS flag on your "optimal taper ratio" theory for the Spiteful. It just does not make any sense.
Nice power point chart BTW.
On the subject of your wooden model spitfire.
This report (http://naca.larc.nasa.gov/reports/1937/naca-report-547/) was written in 1937.
Thegreat advances in monoplane fighter design in Germany made it clear to the RAF that it needed a new home-defence interceptor. Impressed with Mitchell's Type 300, the Air Ministry drew up a specification (F.37/34) around the Type 300 design, and in 1936 ordered 310 production examples of the new fighter. A prototype first flew in March 1936, and the first production Spitfire Mark Is reached No. 19 Squadron RAF at Duxford in August 1938.
Too late for it to applied to a wooden model of the Spitfire. It applies correction for the shape of the tunnel and the model or test aircraft being mounted off-center of the stream. So what was the shape of the tunnel your model was tested in? Mute point because this theory came out AFTER the 1/6th scale model of the spitfire was tested. Therefore no corrections could have been applied.
Crumpp
-
Originally posted by Crumpp
If the physical construction of the wing is not elliptical then you must manipulate the wing structure to achieve an elliptical distribution. This is done by twist. The Spiteful does not have elliptical wing construction. It has NO twist either to move it closer to elliptical.
Well, here is a typical example how you don't understand at all what has been under discussion in this thread. The lift distribution can be manipulated both ways; with taper ratio and the washout. Here is a lift distribution chart from the Glauert book which shows the effect of the optimal taper ratio:
(http://personal.inet.fi/koti/soon.moro/gl.jpg)
Same can be found in the other form (Figure 2-44a) from Perkins&Hage. As Badboy pointed out above with example of the F6F, the tapered (untwisted) wing can reach almost elliptical distribution with optimal taper ratio as can be seen from the chart above.
The Spiteful, Mustang and Fw 190 all had about optimal taper ratio (around the 0,5) but the washout in the wings of the Mustang and Fw 190 moved lift distribution inwards just like washout did same thing in the case of the Spitfire. This can be confirmed from the Lednicer's lift distribution chart:
(http://personal.inet.fi/koti/soon.moro/sp.jpg)
Basicly if you want to prove something else, you should prove that Lednicer is wrong. Besides, this has abolute nothing to do with my calculations.
Originally posted by Crumpp
On the subject of your wooden model spitfire.
This report was written in 1937.
If you had actually read the document you would have found out that it was written 1935 (http://naca.larc.nasa.gov/reports/1937/naca-report-547/index.cgi?page0013.gif) and in the NACA references it is claimed for year 1935. The wind tunnel tests of the Spitfire I have refered are from year 1943 as can be confirmed from the catalogue of the PRO.
gripen
-
The lift distribution can be manipulated both ways; with taper ratio and the washout.
That is funny. For someone who "does not understand" I sure made that my whole point in the first post about the Spiteful.
You don't have a lift distribution chart for the Spiteful, Gripen. So you can't really make any claims about it!
Other than what the math shows.
The Spiteful, Mustang and Fw 190 all had about optimal taper ratio
So tell me how any formula that uses both aspect ratio and taper ratio arrives at result in which the Spiteful beats the FW-190??
Taper ratio's are equal and the FW-190 has a higher Aspect Ratio. Mathmatically it does not make sense Gripen. Unless your manipulating data.
Your whole argument is not logical.
If you had actually read the document you would have found out that it was written 1935 and in the NACA references it is claimed for year 1935. The wind tunnel tests of the Spitfire I have refered are from year 1943 as can be confirmed from the catalogue of the PRO.
AS the paper says Gripen:
Silverstein, Abe White, James A NACA Report 547 1937
http://naca.larc.nasa.gov/reports/1937/naca-report-547/
It was published in 1937.
A year after the prototype Spitfire was flying.
A prototype first flew in March 1936, and the first production Spitfire Mark Is reached No. 19 Squadron RAF at Duxford in August 1938.
The wind tunnel tests of the Spitfire I have refered are from year 1943 as can be confirmed from the catalogue of the PRO.
Double standard here Gripen. You won't take anyone elses word on a document and you expect them to take yours?
If it is from 1943, how can anyone be sure it was a varient that even flew during the war? The Mk XIV was in service in October 1943.
http://www.csd.uwo.ca/~pettypi/elevon/gustin_military/spit/spit1943.html
It was not approved for (+25) boost until much later in the war.
Crumpp
-
Hello Guys.
I've not read the thread to well, but I saw you comparing the 190 wing with the one of the Spiteful.
Now, the Spiteful was supposed to incorporate laminar flow, but with insufficient success as far as I know.
Now the Spiteful was faster than the ultimate Spitfire, however not by so much I think, and also with more power (I think again)
So, just my 5 cents, but take a look at the Spiteful's performance vs say Spit XIV or 21, and compare it with the power.
As far as I remember the Spiteful wing turned out totally as a slight disappointment, - i.e., the original Spitfire wing was actually better than belived to be.
Just from my head, but please have a peek into it ;)
-
Originally posted by Crumpp
For someone who "does not understand" I sure made that my whole point in the first post about the Spiteful.
The Spiteful has taper ratio 0,47 and no washout which yields near elliptical lift distribution. Basicly you did not understand that also taper changes lift distribution.
Originally posted by Crumpp
You don't have a lift distribution chart for the Spiteful, Gripen. So you can't really make any claims about it!
Other than what the math shows.
Glauert corrections are well known and accepted part of the wing theory. Actually the modern definition of the e factor still contains Glauert correction embedded in it as can be seen above.
Besides I'm not making any particular claims or comparisons other than testing if the modern version of the e factor formula works and it seem to work rather well in the case of the plane which had no washout (the effect of the washout is much more difficult to calculate).
Originally posted by Crumpp
So tell me how any formula that uses both aspect ratio and taper ratio arrives at result in which the Spiteful beats the FW-190??
Glauert made corrections for aspect ratio and taper ratio as well as for the washout. Get his book if you are interested to calculate, I have no any particular need to calculate or prove something else than what the drag data shows.
Start a new thread if you want to compare planes this thread is about determination of the e factor not about comparisons.
Originally posted by Crumpp
It was published in 1937.
It was part of the 1937 periodical (starting from p. 135) it was written 1935 and published same year as can be confirmed for example from here (http://naca.larc.nasa.gov/reports/1948/naca-rm-a7i16/index.cgi?page0014.gif). NACA periodicals are large collections of reports which were published individually earlier.
The wind tunnel tests for the Spitfire which I have been refering are from year 1943. The reference number in the PRO is DSIR 23/12576, get the report if you don't believe and the year can be confirmed from the catalogue too.
gripen
-
were you guys in the PRO?
I was there once, and definately want to go again.
It needs some planning if you want to get something out of it though, you may need to place orders in advance etc.
Anyway, I am planning for London, if (economically) possible, this autumn.
If you guys want to meet up, join up, team up, etc, just let me know!!! All for it, and this applies to you all. (even Izzy)
I mean, this is no joke, quite serious. And if you haven't been to London, well, some have (me like 10-20 times), and of course some AH'ers live there. So, anyway, if on the hunt for documents, pictures or data, London IS the place, and I'll be all ready to assist.
Regards.
Angus.
-
Originally posted by Angus
were you guys in the PRO?
I was there once, and definately want to go again.
It needs some planning if you want to get something out of it though, you may need to place orders in advance etc.
I've been there couple times, at least in my case there have been no need to order something in advance. Their catalogue is very usefull. Nowadays they allowe digital cameras in the reading rooms.
gripen
-
[:rolleyes:
-
The Spiteful has taper ratio 0,47 and no washout which yields near elliptical lift distribution. Basicly you did not understand that also taper changes lift distribution.
Proof you did not read my very post on the Spiteful and are simply flamebaiting.
The Spiteful has taper ratio 0,47 and no washout which yields near elliptical lift distribution. Basicly you did not understand that also taper changes lift distribution.
Near Perfect elliptical distribution is .45 according to this report. (http://naca.larc.nasa.gov/reports/1947/naca-tn-1491/index.cgi?page0002.gif)
.47 is too high. Again. Simply post the Lift distribution chart for the Spiteful.
Start a new thread if you want to compare planes this thread is about determination of the e factor not about comparisons.
Facts are Gripen. You can't explain it. It does not work out mathmatically with any formula using Glauerts corrections. The Spiteful should not have an e factor significantly higher than the FW-190.
Taper ratio for the Spitful = .47
Taper ratio for the FW-190 = .44
Aspect Ratio for the Spitful = 5.81
Aspect Ratio for the FW-190 = 6.02
As Glauerts corrections show. Higher Aspect Ratio is more beneficial to wing efficiency and the taper ratios are near equal.
The wind tunnel tests for the Spitfire which I have been refering are from year 1943. The reference number in the PRO is DSIR 23/12576, get the report if you don't believe and the year can be confirmed from the catalogue too.
So it it's not even a varient used in World War II.
Crumpp
-
Originally posted by gripen
Generally the values in the Oswald's chart are probably a bit on high side; the value of e must be always be less than 1 as Wood notes. Practical maximum is probably around 0,95 for the flying wings. Here is some values (http://web.usna.navy.mil/~dfr/flttst1b.pdf) for modern planes from Navy web site:
"The value e can be estimated for a particular model of aircraft by various methods.
One of the simplest is by comparison with aircraft of similar configuration. For most
aircraft e has a value of about 0.6 to 0.8. Some values of e for various aircraft as
measured at Mississippi State University are
Bellancer Crusair (low-wing) 0.55
Learstar (twin engine mid-wing) 0.67
Cessna 170 (high-wing) 0.74
RJ-5 Glider (high-wing) 0.79"
I’m still not happy to use those values in the context of WWII aircraft because the aircraft in the list above are not really intended to operate efficiently under high load factors, they rarely pull more than 3g, while WWII fighters were intended to be more efficient during loaded maneuvers.
Originally posted by gripen
A practical comparison can be done with the Spiteful (no washout) wind tunnel data and calculation for entire airframe using the estimated value of K 0,01 (viscous drag) from Perkins&Hage and Glauert correction factor 0,011 for taper ratio about 0,5 (induced drag):
Wind tunnel => e=0,81
Calculated => e=0,84
Which is a pretty good result given the generalized form of the estimation.
Agreed, and that’s much more like the sort of value I think we should expect for fighters. Also, I would expect the theoretical value to be slightly higher, as they are, because the calculations don’t allow for a number of factors, such as distortion of the wingtip vortices due to viscosity, that reduces the efficiency.
Here are some more comparisons between calculated and experimental values for e shown in the last column on the right.
(http://www.badz.pwp.blueyonder.co.uk/images/ecompare.jpg)
The author of this chart claims to be comparing Oswald’s efficiency factor, but he actually only appears to have tested flying wings. Though I can’t be certain. For example, the elliptical wing in that table has an experimental value of 0.92 and a calculated value of 0.93 and all of the wings with an aspect ratio 6 and taper ratio of 0.5 are very close to an e of 0.9 both experimentally and calculated. You can compensate for the incomplete value of e by including an estimated value for k in the calculation, but the value should be lower than normal because a flying wing would already include some of it in the original value. Doing that would, I anticipate, bring the value back down between 0.8 and 0.85 as before.
Badboy
-
Vmax (IAS) @ 5000’ @ WEP
By Rank Order
By Year
Rank Order – Current Version
Rank
Aircraft
3.05
1
262
461
2
190D9
358
3
Ki84
354
4
J2M2
353
5
109K
353
6
Yak 3
349
7
P47D
349
8
J2M3
346
9
F4u1a
344
10
F4u1D
344
11
Corsair II
344
12
Corsair IV
344
13
F4u4
344
14
P51D
342
15
Spit 14
342
16
P51B
341
17
Fw190A4
339
18
P38J
329
19
P38L
329
20
190A8
326
21
109G6
323
22
C205
322
23
F4u1
321
24
Corsair I
321
25
P47C
321
26
Yak 9
320
27
Ki44-IIb
317
28
P39Q
315
29
P38F
313
30
Mossie NFII
313
31
P400
312
32
Mossie VI
312
33
P39D
308
34
Spit 9
308
35
Ki61-1b
308
36
F6F5
308
37
Ki61-Ic
308
38
C202
306
39
109F4
303
40
109G6R6
303
41
A6M3
295
42
A6M5a
295
43
SPIT 5
294
44
109F1
292
45
Seafire 2
291
46
110G2
290
47
Ki43
288
48
P40E
288
49
P40B
287
50
Spitfire 1
285
51
109E4
284
52
FM2
284
53
A36
279
54
A6M2
271
55
Hurri2c
269
56
F4F3
267
57
F4F4
263
58
Hurricane 1
262
-
Originally posted by Crumpp
Proof you did not read my very post on the Spiteful and are simply flamebaiting.
You said in your post that:
"If the physical construction of the wing is not elliptical then you must manipulate the wing structure to achieve an elliptical distribution. This is done by twist. The Spiteful does not have elliptical wing construction. It has NO twist either to move it closer to elliptical. "
You simply did not understand that the about optimal taper ratio of the Spiteful yields near elliptical lift distribution despite it had no elliptical construction nor washout.
Originally posted by Crumpp
Near Perfect elliptical distribution is .45 according to this report. (http://naca.larc.nasa.gov/reports/1947/naca-tn-1491/index.cgi?page0002.gif)
As can be seen from the Fig. 2-44a in the Perkins&Hage, the taper ratio of the untwisted wing around the 0,4-0,5 yields near elliptical lift distribution, as an example see Badboy's post on taper ratio 0,52 which yielded induced drag factor (u) over 0,99 with computer analysis.
Originally posted by Crumpp
You can't explain it. It does not work out mathmatically with any formula using Glauerts corrections.
If the tapered wing has about optimal taper ratio it also has near elliptical lift distribution. Now if washout is added, it moves lift distribution inwards which yields less optimal lift distribution as can be seen from Lednicer's chart and can be calculated with Glauert corrections. It's up to you to prove that they are wrong.
Originally posted by Crumpp
So it it's not even a varient used in World War II.
It's up to you to prove that the Spitfire I was not used during WWII.
Originally posted by Badboy
I’m still not happy to use those values in the context of WWII aircraft because the aircraft in the list above are not really intended to operate efficiently under high load factors, they rarely pull more than 3g, while WWII fighters were intended to be more efficient during loaded maneuvers.
Neither I'm using those values for the WWII fighters. The drag data I have checked so far indicate that the e factor was typically around 0,75-0,8 in the Cl range say 0,2-1,0 (assuming that there was a linear stage around these Cl values). The problem here is that near Clmax (generally around Cl 1,4-1,8), like in the high g maneuvering near corner speed, the Cd/CL^2 curve is no more linear and it's difficult to say if the e factor defined normal way works well for performance calculations in this kind of maneuvering.
Another problem in the wind tunnel data is that in most cases measurements are made without propeller so slipstream effects are not counted.
Originally posted by Badboy
Agreed, and that’s much more like the sort of value I think we should expect for fighters. Also, I would expect the theoretical value to be slightly higher, as they are, because the calculations don’t allow for a number of factors, such as distortion of the wingtip vortices due to viscosity, that reduces the efficiency.
Yep, very true, I used purely approximated value of the K 0,01. The range for the value of the K (viscous drag multiplier in the e factor formula) given in the Perkins&Hage is 0,009-0,012 which yields e factor range 0,8-0,85 for a plane with about optimal taper ratio and no washout.
gripen
-
You simply did not understand that the about optimal taper ratio of the Spiteful yields near elliptical lift distribution despite it had no elliptical construction nor washout.
Way to manipulate words Gripen!
I simply understand the english adjectives, "Perfect" and "near perfect".
the taper ratio of the untwisted wing around the 0,4-0,5 yields near elliptical lift distribution,
So now it's a range! What happenend to:
The Spiteful has taper ratio 0,47 and no washout which yields near elliptical lift distribution.
You seem very hung up on lift ditribution for the entire wing. While it does effect efficiency, a "perfect ellipse" is not a desirable feature in a practical aircraft, Gripen. The most important part of lift distribution for induced drag is the tip. Designers will strive to get elliptical distribution for the wingtip.
The rest of the lift distribution is usually manipulated to improve the stall and avoid the stall characteristics of a perfect ellipse.
The other problem with elliptical wings is the stall characteristics. It is much safer to design an airplane so that the wing stalls first at the root, leaving the outer portion of the wing, (where the ailerons are) still flying. An elliptical wing however, will tend to stall uniformly all along the span (see the diagram below.) The "fix" for this situation is washout, but that will reduce the theoretical gains in induced drag. Therefore, we are unlikely to see a great resurgence in the use of elliptical wings, except in situations where appearance dictates.
http://142.26.194.131/aerodynamics1/Drag/Page8.html
The wing-tip shape, being at the point where the tip vortices are produced, appears to be of more importance in minimizing tip vortex formation and thus minimizing induced drag.
Taper and twist are perhaps of greater importance in dealing with the problem of stalling.
http://www.centennialofflight.gov/essay/Theories_of_Flight/Reducing_Induced_Drag/TH16.htm
So in effect you cannot explain why your math gives the FW-190 a much lower e factor than the Spitfeful.
You keep avoiding the question of how that is possible with the aircraft having near perfect taper ratios and the FW-190 having a significant advantage in Aspect Ratio?
As for your wooden model, now I am confused about it. First it was a 1943 report so that it could have used an NACA report calculating corrections of off center flow around the tail which included formulas that brought the model closer to the actual aircraft published in 1937.
Gripen says:
The wind tunnel tests of the Spitfire I have refered are from year 1943 as can be confirmed from the catalogue of the PRO.
Now it is a Spitfire I?
Gripen says:
It's up to you to prove that the Spitfire I was not used during WWII.
Which is it?
Great chart Badboy. Interesting the wide range of shapes with very similar efficiency.
Crumpp
-
Originally posted by Crumpp
Way to manipulate words Gripen!
Semantics are not subject of this thread. The Spiteful had near elipptical lift distribution as can be proofed with Glauert corrections. Try to live with that.
Originally posted by Crumpp
You keep avoiding the question of how that is possible with the aircraft having near perfect taper ratios and the FW-190 having a significant advantage in Aspect Ratio?
Allready answered above:
"If the tapered wing has about optimal taper ratio it also has near elliptical lift distribution. Now if washout is added, it moves lift distribution inwards which yields less optimal lift distribution as can be seen from Lednicer's chart and can be calculated with Glauert corrections."
Besides the difference between 0,78 and 0,81 is small just like the difference between 0,8 and 0,85.
Originally posted by Crumpp
As for your wooden model, now I am confused about it. First it was a 1943 report so that it could have used an NACA report calculating corrections of off center flow around the tail which included formulas that brought the model closer to the actual aircraft published in 1937.
Well, you are certainly confused, it is a 1943 report as can be confirmed from the catalogue of the PRO.
Originally posted by Crumpp
Now it is a Spitfire I?
It is a Spitfire I as can be confirmed from the catalogue of the PRO.
gripen
-
Originally posted by gripen
Neither I'm using those values for the WWII fighters. The drag data I have checked so far indicate that the e factor was typically around 0,75-0,8 in the Cl range say 0,2-1,0 (assuming that there was a linear stage around these Cl values). The problem here is that near Clmax (generally around Cl 1,4-1,8), like in the high g maneuvering near corner speed, the Cd/CL^2 curve is no more linear and it's difficult to say if the e factor defined normal way works well for performance calculations in this kind of maneuvering.
gripen
The efficiency factor does work during that kind of maneuvering, what doesn’t work so well is the parabolic drag polar. The fact is that all the theory we have been discussing has evolved around the assumption that the relationships can be fitted with a parabolic polar. That turns out to be a very reasonable assumption for aircraft that have low AoA limits and depend exclusively on conventional lift. However, it is only approximately true, and in situations such as when other forms of lift are involved, or when the flow becomes turbulent near the edge of the envelope, it becomes more difficult to ignore. In those situations, it isn’t the value of e that needs to be modified, something other than a parabolic polar should be used.
The popularity of parabolic polars, and quadratic relationships and their appeal in aerodynamics stems from the fact that most of the resulting equations can be solved analytically, and generally they do produce good results, and errors are smoothed out to some extent within the coefficients. However, when you work with real aircraft, it isn’t always possible to use those simple relationships, and there are a number of alternatives. Firstly, you can model with non-quadratic relationships, but then the equations do not always have analytical solutions, (but they sometimes do) which doesn’t really matter because you can solve them using numerical methods, it just requires more computer time. There are other ways to match the real curves, very accurately, or at least with any degree of fidelity you wish, but it involves working with large quantities of tabular data. Both of those methods entirely resolve the issues of partial fit discrepancies for parabolic curves and enable very close predictions, even at the extremes of the envelope. I use those methods when ever I work with real aircraft, for EM analysis for example, but if you are analyzing what is happening in a simulation, such as AH for example, the traditional method is fine, because that is what is being used in the flight model anyway.
Hope that helps…
Badboy
-
Semantics are not subject of this thread. The Spiteful had near elipptical lift distribution as can be proofed with Glauert corrections. Try to live with that.
Your making semantics the subject by accusing other of not understanding what you your self do not understand.
"If the tapered wing has about optimal taper ratio it also has near elliptical lift distribution. Now if washout is added, it moves lift distribution inwards which yields less optimal lift distribution as can be seen from Lednicer's chart and can be calculated with Glauert corrections."
Certainly to control the less desirable stall characteristics of perfect elliptical lift distribution.
Your attempting to say that that efficiency was completely destroyed down to .78?? Come on, Gripen.
All the working aeronautical engineers looking at the same data you have already say it is .87. Much more reasonable given the Taper Ratio and the Aspect Ratio. The FW-190's wing as listed on Badboys chart with NO twist yields an efficiency of .91.
Aeronautical engineers at the time of the FW-190's design were well aware of the benefits of elliptical wingtip distribution and how to attain it.
It is easy to see from this lift distribution chart that the FW-190 and the P51 were closer to elliptical distribution at the wingtips than the Spitfires.
(http://www.onpoi.net/ah/pics/users/503_1105299227_span.jpg)
The wingtips being the most important part of wing efficiency, this would go a long way towards correcting the rest of the elliptical departure in the wing structure. This is shown by every accepted formula in aeronautics.
Your value of .78 is calculated from your half-baked "Lift distribution" calculations and has no place in the real world.
I don't believe for an instant you even calculated off that FW-190 chart.
Especially when compared to your results from the Spitefuls wing. Taper ratio are within .03 between the two and the FW-190 has a decisive Aspect Ratio advantage of .21. You come up with a .84 result for the Spiteful! Poppycock on your FW-190 results. Your manipulating data in an attempt for recover from the "Draining E turns Thread".
Crumpp
-
Originally posted by Badboy
The efficiency factor does work during that kind of maneuvering, what doesn’t work so well is the parabolic drag polar.
I can't really follow now, a constant value of the e results perfectly parabolic drag polar. As an example we can assume a plane with rectangular wing with AR 6 and e factor 0,75, the formula for the drag rise written with e factor is then:
Cd = Cd0 + Cl^2/(pi*6*0,75)
Same can be also written in the form of the simple polar as:
Cd = Cd0 + 0,07073553*Cl^2
It can be also written in the form of the e factor formula (Glauert correction factor 0,05):
Cd = Cd0 + 0,0150313*Cl^2 + Cl^2/(pi*6)*(1+0,05)
All these three ways to calculate drag rise results exactly similar perfectly parabolic drag polar which does not work particularly well near Clmax in most cases.
Originally posted by Crumpp
Certainly to control the less desirable stall characteristics of perfect elliptical lift distribution.
This thread is not about the stalling characters.
Originally posted by Crumpp
Your attempting to say that that efficiency was completely destroyed down to .78??
No one says that the efficiency was completely destroyded, 0,78 is typical value for the WWII fighter.
Originally posted by Crumpp
All the working aeronautical engineers looking at the same data you have already say it is .87.
Who? There is other answer than 0,78 for the value of the K 1,24 in the Fw chart.
Originally posted by Crumpp
Much more reasonable given the Taper Ratio and the Aspect Ratio. The FW-190's wing as listed on Badboys chart with NO twist yields an efficiency of .91.
There is no Fw 190 wing in that chart.
Originally posted by Crumpp
It is easy to see from this lift distribution chart that the FW-190 and the P51 were closer to elliptical distribution at the wingtips than the Spitfires.
Lednicer about the lift distribution of the Spitfire:
"... loading distribution is not elliptical, though it is probably most optimum of three from the induced drag standpoint."
Originally posted by Crumpp
You come up with a .84 result for the Spiteful!
Nonsense, the wind tunnel data results 0,81, the value 0,84 is just a test result to check validity of the formula with approximated value of K.
gripen
-
This thread is not about the stalling characters.
Again you don't read. My point was not about stalling, Gripen. It was the FACT that lift distribution beyond the wingtip has little effect on efficiency, just like the aeronautical engineers say.
There is no Fw 190 wing in that chart.
The NACA 230015-009 is the FW-190 Wing, Gripen, minus 2 degrees of twist.
"... loading distribution is not elliptical, though it is probably most optimum of three from the induced drag standpoint."
Exactly. Once again you completely miss the mark on the point. You have a talent for completely ignoring data which does not suite your vision and twisting the words of others.
Crumpp says:
It is easy to see from this lift distribution chart that the FW-190 and the P51 were closer to elliptical distribution at the wingtips than the Spitfires.
Completely different meaning that what your trying to twist out of it.
Since wingtip elliptical distribution is the MOST important part of induced drag reduction, it's not hard to see why the Spitfire was not that far ahead of either the P51 or the FW-190. No matter what formula the spread remains the same between the three.
Just check out the elliptical planform wing 4412 on Badboys chart. It's e factor is .92 and the 230015-009 is .91.
Crumpp
-
Originally posted by Crumpp
Again you don't read. My point was not about stalling, Gripen. It was the FACT that lift distribution beyond the wingtip has little effect on efficiency, just like the aeronautical engineers say.
The lift distribution is function of the entire wing as pointed out by Prandtl, Glauert and entire wing theory. Try to live with that.
Originally posted by Crumpp
The NACA 230015-009 is the FW-190 Wing, Gripen, minus 2 degrees of twist.
Nonsense, this wing just has same profile but not same characters (washout, shape etc.).
Originally posted by Crumpp
You have a talent for completely ignoring data which does not suite your vision and twisting the words of others.
You are the one who can't accept what Lednicer says.
Originally posted by Crumpp
Just check out the elliptical planform wing 4412 on Badboys chart. It's e factor is .92 and the 230015-009 is .91.
The 4412 is a wing profile and again it has no washout. Besides the source of the chart is not documented and we are talking about the e factor of the entire airframe not just wing.
gripen
-
Besides the source of the chart is not documented and we are talking about the e factor of the entire airframe not just wing.
Oh now the chart is bad, huh?
The lift distribution is function of the entire wing as pointed out by Prandtl, Glauert and entire wing theory. Try to live with that.
Again you twist what others say.
NASA Says:
The wing-tip shape, being at the point of production of the tip vortices, appears to be of more importance in minimizing tip vortex formation and thus minimizing induced drag.
http://history.nasa.gov/SP-367/chapt4.htm#f62
NASA says:
Taper and twist are perhaps of greater importance when the problem of stalling is discussed later.
http://history.nasa.gov/SP-367/chapt4.htm#f62
Is the light bulb on Gripen?
Nonsense, this wing just has same profile but not same characters (washout, shape etc.).
Afraid it is Gripen. As I pointed out earlier, it the FW-190 wing without the 2 degree twist to improve the stall.
Crumpp
-
Originally posted by Crumpp
Oh now the chart is bad, huh?
Generally if some one brings in the data the source should be noted. In this case I found the source (http://naca.larc.nasa.gov/reports/1938/naca-report-627/) myself.
Actually it confirms what I said about the e factor near Clmax as well as gives the reason for e value of the 4412 wing:
"This curve is typical for the wings and shows how the efficiency-factor curve departs from the Cd curves below Cl=0.2 to 0,4 and above Cl=1.0. Reference to table I shows that that NACA 24-30-8.50 and 2R1-15-8.50 wings which have the largest Clopt, have values of e equal to and larger than e, respectively, for the elliptical NACA 4412 wing. This result is obtained because shifting the Cde curve to the right makes it fit a flatter e curve, and hence one with a higher value of e. If Clop, had been zero for all the wings and they had differed only in plan form, the values of e would indicate the departure of the drag of the wings from that of the ideal elliptical wing. The wings, in fact are sufficiently similar and the variations of the Cd0 values with lift are near enough alike so that there is a general reduction of e as the wings depart from the elliptical plan form toward the wings of high taper."
Originally posted by Crumpp
Blaah Blaah...
The situation is very simple here; your theories against Lednicer's calculations and word, data from Fw and RAE as well theory from Prandtl and Glauert. Entire wing generate vortices and very small differences near wing tip are neglible as can be proofed from the Glauert's book, note that half tapered wing has nearly same e as elliptical despite differences in the lift distribution:
(http://personal.inet.fi/koti/soon.moro/gl.jpg)
Besides no one else than you is comparing planes here, all the data here is just to show what was the normal range of e for WWII. If you don't like the data, it's not my problem nor subject of this thread.
gripen
-
The situation is very simple here; your theories against Lednicer's calculations and word, data from Fw and RAE as well theory from Prandtl and Glauert.
That's funny! It's not my theories, its the exact folks your quoting above backed up by NASA.
Nice attempt at a bait and switch.
Actually it confirms what I said about the e factor near Clmax as well as gives the reason for e value of the 4412 wing:
Which has nothing to do with our discussion.
Generally if some one brings in the data the source should be noted. In this case I found the source myself.
So now it is good. Make sure you tell Badboy about the source standards. Is that the same standard as your Spitfire wooden model report?
Actually it confirms what I said about the e factor near Clmax as well as gives the reason for e value of the 4412 wing:
No one is arguing this Gripen. Your a master of the obvious and seriously attempting a straw man rebuttal.
Gripen says:
Entire wing generate vortices and very small differences near wing tip are neglible as can be proofed from the Glauert's book, note that half tapered wing has nearly same e as elliptical despite differences in the lift distribution:
You need to inform NASA cause they are not on the same sheet of music. It's the tip vortices that are the main contributor to induced drag.
NASA Says:
The wing-tip shape, being at the point of production of the tip vortices, appears to be of more importance in minimizing tip vortex formation and thus minimizing induced drag.
http://history.nasa.gov/SP-367/chapt4.htm#f61
Sure your not making up another Gripen Theory of Aerodynamics?
Crumpp
-
Originally posted by Crumpp
It's not my theories, its the exact folks your quoting above backed up by NASA.
The difference in 1/10 part of the wing in the Lednicers lift distribution chart is much smaller than the difference in the Glauert chart which did no difference. NASA site claims nothing about small differences in the lift distribution chart, but Lednicer's opinion on his data is known as well as backed up by data from Fw and RAE and also supported by theory.
Besides, according to Lednicer's chart Mustang does the smallest tip vortices and Fw the largest of the three.
Originally posted by Crumpp
Which has nothing to do with our discussion.
Basicly it says that to compare the plan form characters of the wing all other characters (like profile) should be equall, that was not case when you compared other wings to elliptical wing.
Originally posted by Crumpp
Make sure you tell Badboy about the source standards. Is that the same standard as your Spitfire wooden model report?
It's allways a good standard to identify sources and I have given the source of the report on Spitfire I.
gripen
-
The difference in 1/10 part of the wing in the Lednicers lift distribution chart is much smaller than the difference in the Glauert chart which did no difference.
So what Gripen? The difference is the most important part of lift distribution for induced drag production. Think of it like wing loading is to turn performance. It is just one of the factors. However wing loading is the most important factor in turn performance and you can get a feel for turn performance looking at it.
1/10th is kind of an exaggeration too. 20 percent at the tip Gripen and that is not counting the other 20 percent at the root were the FW-190 is equal or better.
but Lednicer's opinion on his data is known as well as backed up by data from Fw and RAEand also supported by theory.
Only problem with any of this is YOUR THEORY. You need to fit your world to the facts and not the facts to your world.
Lednicer's opinion is published in his article that you seem to read much more into than what he says about efficiency factor differences. His main point is the Spitfire does not have the elliptical lift distribution commonly attributed to it.
(http://www.onpoi.net/ah/pics/users/503_1105387783_spitfireelliptical.jpg)
So based on the calculations and the chart, the Spitfire does have a better efficiency. Because it's wingtip lift distribution is the farthest from elliptical due to its uniform twist, though it's advantage is not huge. In fact is close enough that Lednicer refrains from calling it a fact that the Spitfire has the advantage.
Just as NASA says. Wingtip lift distribution is the main contributing factor to induced drag formation. The rest of the lift distribution is a minor contributor and commonly the wing design is manipulated AWAY from elliptical to improve the stall.
Crumpp
-
Originally posted by gripen
All these three ways to calculate drag rise results exactly similar perfectly parabolic drag polar which does not work particularly well near Clmax in most cases.
Exactly, and that’s as good as it gets, the situation actually gets worse for many other wing configurations and the departure from a parabolic polar increases. However, what is failing is not the efficiency factor, but the parabolic drag polar. It was only ever approximately true, and in situations where other forms of lift are involved, or near the edge of the envelope, the parabolic relationship can no longer be used. When you said:
Originally posted by gripen
The problem here is that near Clmax (generally around Cl 1,4-1,8), like in the high g maneuvering near corner speed, the Cd/CL^2 curve is no more linear and it's difficult to say if the e factor defined normal way works well for performance calculations in this kind of maneuvering.
gripen
It sounded to me as though you were thinking that it was the e factor that might need to be defined differently, when it is the parabolic relationship at fault. There are other methods to match the real curves with any degree of fidelity you wish, but not with quadratics. Those methods don’t suffer from the issues of partial fit discrepancies (that occur because parabolic curves are only good for non-turbulent flow and traditional lift) and are able to provide very close predictions, even at the extremes of the envelope.
However, parabolic polars, and quadratic relationships are popular, and their appeal in aerodynamics stems from the fact that most of the resulting equations can be solved analytically, and can be manipulated easily in order to learn about the relationships involved. Also, they do generally produce good results, and errors are smoothed out to some extent within the coefficients, so that they are satisfactory for most practical purposes.
All I’m saying is that when that relationship breaks down, as it does near the extremes of the envelope and even more widely in many other wing configurations, you simply can’t rely on quadratic equations any more, and need to use other methods.
Hope that helps…
Badboy
-
Originally posted by gripen
Generally if some one brings in the data the source should be noted. In this case I found the source (http://naca.larc.nasa.gov/reports/1938/naca-report-627/) myself.
gripen
My apologies… I thought I had posted the link.
Did you notice the author’s comments that he was specifically trying to find an efficiency factor comparable to the airplane efficiency factor, and he then provides a reference to Oswald’s paper to support his point. But he then appears to have tested wings without the fuselage, I guess folk have been confused on that point for a long time. I’ve seen it done both ways almost indiscriminately for many years :)
Badboy
PS
But I don't feel too bad about forgetting the source, after all I've given you some very good ones ;)
Edited in the PS
-
OMG, this thread is getting so big, that it is actually THREATENING the Spit-Messer thread :D
Anyway, I saw you mentioning the wing tip configurations as ell as the Spitfire wing.
Well, for what it's worth, AFAIK, clipping that wing would increase roll rate incredibly, speed slightly, climb would fall, zoom would rise, and turning would drop by some margin, stall would be more vicious.
Would it fit? I'd think yes ...
-
Originally posted by Crumpp
The difference is the most important part of lift distribution for induced drag production.
The tip vortices are the result of the pressure difference between top and bottom of the wing and therefore a wing with lot of lift near tip produce larger tip vortices than a wing with less lift near tip. The elliptical lift distribution means constant downwash along the wingspan. Funny thing in your theory is that if your favorite plane has more lift (downwash) near tip than some others, it also produces larger tip vortices. In addition it should be noted that if there is no constant downwash in the rest of the wing, there is no such in the tip. Generally a wing with least variation of downwash along the span produces least induced drag.
Originally posted by Badboy
However, what is failing is not the efficiency factor, but the parabolic drag polar. It was only ever approximately true, and in situations where other forms of lift are involved, or near the edge of the envelope, the parabolic relationship can no longer be used.
The good point in the parabolic drag polar is that it has a theoretical explanation in the lifting line theory. I could form a formula, say the form Cd=Cd0+X1*Cl^3+X2*Cl^2 (X1 and X2 being constants) or something, which might work better at low and high Cl values, but such formula has no theoretical background.
Anyway, I've been under impression that in the AH the drag rise is modeled with AoA ie not by making some kind of Cl based model.
Originally posted by Badboy
Did you notice the author’s comments that he was specifically trying to find an efficiency factor comparable to the airplane efficiency factor, and he then provides a reference to Oswald’s paper to support his point. But he then appears to have tested wings without the fuselage, I guess folk have been confused on that point for a long time. I’ve seen it done both ways almost indiscriminately for many years :)
Yep, generally the e factor for the wing only can be estimated good enough but muddy waters start when estimating the rest of the airfame so. Wood gave one solution but it's difficult to say if it works generally well. RAE wind tunnel data for the Mustang I contains drag data for the wing only which result value of e about 0,88 for the wing, while the entire airframe has value of e about 0,77. Other planes were not tested with wing only at higher Cl range.
Angus,
Start a new thread if you want to compare planes.
gripen
-
Originally posted by gripen
The good point in the parabolic drag polar is that it has a theoretical explanation in the lifting line theory. I could form a formula, say the form Cd=Cd0+X1*Cl^3+X2*Cl^2 (X1 and X2 being constants) or something, which might work better at low and high Cl values, but such formula has no theoretical background.
Yep, but take a look at these three drag polars.
(http://www.badz.pwp.blueyonder.co.uk/images/polarcompare.jpg)
You can see by inspection, that only the NACA 0006 can be modeled with a parabolic polar. The NACA 4412 might yield to curve fitting, and because it has two distinct curved portions with a portion of high curvature between, there is a good chance that the slope of dCd/dCl^2 would have two distinct straight line portions, so it could be treated with two e values, one for each part of the envelope (check your copy of Pope’s book). However, the laminar flow wing with its classic drag bucket won’t yield to either approach. In those cases, a parabolic model is only good for one out of three of them, so a parabolic model may have a good basis in theory, but it is more useful in the classroom than it is in practical aerodynamics. The good news is that all three of those polars can be modeled in exactly the same way by aero engineers working directly from the raw data, and each can be represented with almost any degree of fidelity required, despite the discontinuities.
Originally posted by gripen
Anyway, I've been under impression that in the AH the drag rise is modeled with AoA ie not by making some kind of Cl based model.
I agree, AH has an excellent flight model, that responds well to classical analysis.
Hope that helps...
Badboy
-
Originally posted by Badboy
You can see by inspection, that only the NACA 0006 can be modeled with a parabolic polar. The NACA 4412 might yield to curve fitting, and because it has two distinct curved portions with a portion of high curvature between, there is a good chance that the slope of dCd/dCl^2 would have two distinct straight line portions, so it could be treated with two e values, one for each part of the envelope (check your copy of Pope’s book). However, the laminar flow wing with its classic drag bucket won’t yield to either approach. In those cases, a parabolic model is only good for one out of three of them, so a parabolic model may have a good basis in theory, but it is more useful in the classroom than it is in practical aerodynamics.
True, but section data contains just variation of the profile drag in 2D. If we go 3D and ad induced drag, the polar of the laminar flow wing look much more familiar because the parabolic rise of the induced drag is much stronger than the rise of the section drag (from Abbott&vonDoenhoff):
(http://personal.inet.fi/koti/soon.moro/sect.jpg)
Parabolic estimation can be used but accuracy is not best possible because the shape of the profile drag rise shows a bit in the Cd/Cl^2 curve.
gripen
-
Originally posted by gripen
Parabolic estimation can be used but accuracy is not best possible because the shape of the profile drag rise shows a bit in the Cd/Cl^2 curve.
gripen
Agreed.
Badboy
-
Gripen says:
Funny thing in your theory is...blah blah blah
Nice bait and switch.
Again Gripen it is NASA you are saying it is wrong not me.
Don't try and kill the messenger just because you do not like the message.
The wing-tip shape, being at the point of production of the tip vortices, appears to be of more importance in minimizing tip vortex formation and thus minimizing induced drag.
http://history.nasa.gov/SP-367/chapt4.htm#f61
Crumpp
-
Gripen says:
So I quess this claim is again in the same category as the "wet lifting area" ie you are talking about a thing which does not exist.
David Lednicer was nice enough to answer this question:
Sorry for the delay - I just got back from a week's vacation.
Usually, the wing area of an aircraft includes the part of the wing buried
inside the fuselage. "Wetted Lifting Area" refers to just the part of the
wing that is exposed.
-Dave
Which is exactly how I was using it.
Crumpp
-
there are now more posts in this thread than data points...
-blogs
-
So, netto wing area, so to speak of.
-
(http://personal.inet.fi/koti/temp0/fwpolar.jpg)
For everyone's viewing pleasure here is couple Fw 190 drag polars measured in the Chalais-Meudon. Earlier measurement (from the FW report 6006) results e factor value about 0,69 at Cl range 0,2-1 and later measurement (from the report UM21 1) results e factor value about 0,55 at same Cl range. Early measurements by the Germans in the Chalais-Meudon had a systematical error due to wrong correction factors.
gripen
-
You might want to figure out how to read the plot Gripen.
Keep it up.
All the best,
Crumpp
-
Dear Crumpp,
I wonder what you mean, the report actually contains the Ca and Cw values in the written form so there is no need to read plot.
Regarding the question if the the Ta 152H was tested in the Chalais-Meudon, the chart below is quite self explaining.
gripen
(http://personal.inet.fi/koti/temp0/wt.jpg)
-
Earlier measurement (from the FW report 6006) results e factor value about 0,69 at Cl range 0,2-1 and later measurement (from the report UM21 1) results e factor value about 0,55 at same Cl range. Early measurements by the Germans in the Chalais-Meudon had a systematical error due to wrong correction factors.
Gripen facts are there is nothing wrong with "other plot".
Why don't you post the aircraft set up?
What you have here is the investigation into the FW-190G and ways to reduce it's drag of the wing mounted Zusatzkraftstoffbehalter's..
Of course in your mind the Germans would have been unable to perform a calculation yet in mid thirties the Supermarine was able to a use 1/6th scale wooden model to 100 percent accurately glean 1 for 1 comparable data with a production model Spitfire.....
Secondly how are your figuring your "e" factor out. As we have seen from this thread there is a wide variation for e factor based on the formula and your scale used.
All the best,
Crumpp
-
Dear Crummp,
From the report:
"Der Schlechtere Verlauf der Polaren kann von der Korrektur infolge der Strahlneigung herrühren, die nach früheren Messungen mit 1,5deg angegeben war, nach neueren Messungen aber 0,25deg beträgt."
Basicly all early measurements by the Germans in the Chalais-Meudon were systematically wrong due to this error.
Regarding the condition of the plane see the picture below. In both cases the measured plane was in the clean condition.
The system to calculate e from the polar (or what ever Cd/Cl data) is described several times above. It takes just couple minutes if you understand what you are doing.
gripen
(http://personal.inet.fi/koti/temp0/pic.jpg)
-
You need to get a good translation of that passage Gripen!
Keep trying though....
Wonder what those racks are that the supports are attached too under the wings???
All the best,
Crumpp
-
Oh well, the report is quite clear and it also gives a description how the Germans discovered the error:
"Bei der Auswertung der Widerstandsbeizahlen wurde die Strahlneigung die nach neueren Messungen eines symmetrischen Profils in Normal- und Rückenlage 0,25deg nach unten beträgt, berücksichtigt."
Regarding the condition of the plane the picture is clear (original is even more clear); there is no other supports than the test stand which is the same as used in the earlier measurements. Besides the polars actually indicate that the plane in the later measurement was in the cleaner condition than in the early measurements; at Cd0 the drag is a bit lower. Therefore it's safe to say that differences between the polars are mainly caused by the different correction factors just like the report says.
gripen
-
Gripen the report says the lower values are wrong due to the error.
Think about what the error is Gripen. What do you think a .25 degree AoA will do to your induced drag??
there is no other supports than the test stand which is the same as used in the earlier measurements.
It's not the supports, its the racks under the wing. The Focke-Wulf tragers are clearly visible in the photograph, Gripen.
Peformance differences between an FW-190G and FW-190A:
(http://www.onpoi.net/ah/pics/users/503_1109451277_fw190_g3_s.gif)
As for your contention about the Ta -152H being a calculation. In multiple plane report experimental A/C sometimes are shown as calculated curves. It does not mean that every A/C on the report is a calculation. For example several curves on this report are calculations. Many of them are not:
(http://www.onpoi.net/ah/pics/users/503_1110500796_doraspeed3.3.jpg)
All the best,
Crumpp
-
Dear Crumpp,
The report says clearly that the correction factors for the earlier measurement were wrong. The difference can be easily seen from the polars, in the earlier measurements the Cd0 is higher but at high Cl the Cd values are lower. In the later meassurements Cd0 is lower but the Cd values at high Cl are higher.
Very simple and very clear, try to live with it.
gripen
-
The difference can be easily seen from the polars, in the earlier measurements the Cd0 is higher but at high Cl the Cd values are lower. In the later meassurements Cd0 is lower but the Cd values at high Cl are higher.
Exactly what I just said Gripen.
Earlier larger values of drag are wrong IN THIS REPORT.
As stated in the report it has nothing to do with earlier tests.
Might be worth it to get a translated copy.
However it is much more useful to have one on the FW-190A as opposed to the FW-190G.
Look for the reports on the FW-190A8 prototypes as a helpful hint.
Thanks for the amusement though! I see you have laid awake at night worrying about this discussion.
Have you come across a report on production Spitfires yet? Maybe plots from a real plane?
All the best,
Crumpp
-
Soo... this is long, but which one was faster, the Mk IX or the FW 190A? Low altitude, high altitude?
-
Originally posted by Crumpp
Earlier larger values of drag are wrong IN THIS REPORT.
Oh well, yestarday Crumpp announced:
"Gripen facts are there is nothing wrong with "other plot"."
Besides the report says directly:
"An der gleichen Maschine waren früher schon einmal Sechskomponenten- und Druckverteilungsmessungen durchg eführt; Bericht Nr. 06006 "Messungen an einer Fw 190 im groben Windkanal von Chalais Meudon bei Paris""
Basicly it is the very same plane in the same condition.
Shortly, Crumpp is creating his usual nonsense here.
Kürfurst can open a new thread if he is interested about plane comparisons.
gripen
-
I think Crumpp is making some good points here. I was just asking about level speed, as it seems the 190 was faster at the same power than the Spit. Wouldn`t that point to that it was also more aerodynamic (same power, more speed -> less drag).?
Sorry if I interrupted this classy conversation.
-
Dear Kurfürst,
This thread is about e factor not about comparing planes. Start a new thread if you want such comparison
Regarding "some good points here"; is it a "good point" to change an argument to exactly opposite in one day?
gripen
-
"good point" to change an argument to exactly opposite in one day?
No argument has changed Gripen. Facts are you do not have the information you think you do.
An der gleichen Maschine waren früher schon einmal Sechskomponenten- und Druckverteilungsmessungen durchg eführt; Bericht Nr. 06006 "Messungen an einer Fw 190 im groben Windkanal von Chalais Meudon bei Paris""
All this says Gripen is that they used the same test machine for first six tests of the FW-190G. There are dozens and dozens of windtunnel test's on the FW-190A.
So far:
You have tried to claim a mistake in the windtunnel set up caught and corrected during testing invalidates every test conducted but this one!
As stated in the report the UNCORRECTED test's show HIGHER drag values!
Do you think that higher total drag = higher induced drag = less efficient "e" factor? Did you miss the whole conversation with Badboy?
You have tried to pass an FW-190G off as an FW-190A. Totally different aircraft with very different drag characterisitics.
However, using windtunnel data gleaned from a 1/6th scale wooden model of the Spitfire is perfectly acceptable to mirror a production model in your mind!
All the best,
Crumpp
-
Originally posted by Crumpp
No argument has changed Gripen.
Nonsense, the simple fact is that Crumpp changed his argument exactly opposite in less than 24 hours.
Originally posted by Crumpp
Facts are you do not have the information you think you do.
Well, here we have Crumpp's word against verifyable documentation (I have given the source information).
Originally posted by Crumpp
All this says Gripen is that they used the same test machine for first six tests of the FW-190G.
Ah, Crumpp is changing his argument again. The tested plane is a clean Fw 190A (without pylons or what ever) as can be verified from the pictures.
Originally posted by Crumpp
You have tried to claim a mistake in the windtunnel set up caught and corrected during testing invalidates every test conducted but this one!
The jet inclination error presented above is well documented, basicly all early tests by the Germans had this problem.
Originally posted by Crumpp
As stated in the report the UNCORRECTED test's show HIGHER drag values!
Plain nonsense, the report is clear on this; correction factors for the early measurements were simply wrong.
Originally posted by Crumpp
Do you think that higher total drag = higher induced drag = less efficient "e" factor? Did you miss the whole conversation with Badboy?
Well, I have concensus with Badboy but Crumpp still can't understand the relative concept of the e factor.
gripen
-
Originally posted by Crumpp
...
All this says Gripen is that they used the same test machine for first six tests of the FW-190G.
Crumpp,
"Sechskomponenten" means that all six force components are measured: 3 orthogonal forces (normal, axial, and side forces) and 3 orthogonal moments (rolling, pitching, and yawing moments).
-
h, Crumpp is changing his argument again. The tested plane is a clean Fw 190A (without pylons or what ever) as can be verified from the pictures.
Look at the picture Gripen.
The racks are clearly visible. You need to read my post instead of make up what you wish the world to be.
Well, here we have Crumpp's word against verifyable documentation (I have given the source information).
You need to go back through this thread. I always back up what I say with documentation. Or do you forget the "filled and polished" episode for example?
Facts are Gripen I am much farther along in my research than I was months ago when this whole thread started. Actual aeronautical engineers, not folks like you with a tenious grasp of the theory and a obvious game agenda, have done the analysis from multiple sources. More than one engineer has been nice to enough to lend me a hand with this project.
Well, I have concensus with Badboy
Again. Not how everyone else would see this but if you need to tell yourself that, ok. Badboy taught you about "e" factor. You were not a colleague; you were a student in that conversation.
Plain nonsense, the report is clear on this; correction factors for the early measurements were simply wrong.
In the context of the report Gripen, it refers to earlier test's in that report. That is clearly stated in the report.
Did you stop to think how wind tunnels are operated? Do you think they just flip the switch off, turn the lights off, and go home until the next test? Like any machine they require constant maintenance.
Do you think alarm bells were going off in the Focke Wulf engineering department when the tunnels started returning values that were not even close to the design targets?
The fixing of the wind tunnel IMPROVED the drag over the first few test conducted.
Or do you think it all just happened in a vacuum, Gripen, and you discovered their flawed analysis. Silly Germans, can't do anything right...
All the best,
Crumpp
-
Originally posted by Crumpp
Look at the picture Gripen.
Let's look:
(http://personal.inet.fi/koti/temp0/pic.jpg)
So where are the racks? All we see is parts of the support system. Original picture shows the bottom of the left wing clearly; no racks, pylons or what so ever.
Originally posted by Crumpp
You need to go back through this thread. I always back up what I say with documentation.
Well, Crumpp has changed his stories several times during this thread.
Originally posted by Crumpp
Again. Not how everyone else would see this but if you need to tell yourself that, ok. Badboy taught you about "e" factor. You were not a colleague; you were a student in that conversation.
Anyone can look back this thread and find out who was teaching in the case of the so called generic formulas (which actually were not as has been pointed out).
Originally posted by Crumpp
In the context of the report Gripen, it refers to earlier test's in that report. That is clearly stated in the report.
Nonsense, the report simply quotes earlier results from the Fw report and points out that the differences are mainly caused by the wrong correction factors used before the error was discovered while testing symmetrical wing profiles.
gripen
-
Gripen,
The Focke Wulf trager are clearly visible in the photo.
FW Trager:
(http://www.onpoi.net/ah/pics/users/503_1112398117_picfwtrager.jpg)
Photo:
(http://www.onpoi.net/ah/pics/users/503_1112399313_pic.jpg)
You can clearly see both support rails. Especially if you use a photo editor program to enlarge the image. The rails are labeled 1 and 2 and the arrow points out the back of the fairing.
(http://www.onpoi.net/ah/pics/users/503_1112398293_fwtrager1.jpg)
The only nonsense here is you trying to pass this off as representative of the FW-190A.
All the best,
Crumpp
-
Well, Crumpp creates simply fiction, there is no rack nor rails as can be seen from the profile of the right wing, just some kind of cable coming from the support stand (probably unfinished testing equipment assembly). Funny thing is that in the original picture the wire of the lifter in backround is visible through the supposed "trager".
Besides if there is a "trager" it should be closer fuselage. The Document gives the location of the support stand.
gripen
-
Gripen,
You can see the other one as well. Try enlarging the picture.
Gripen says:
Besides if there is a "trager"
They are both in the picture and in the correct location, Gripen.
Focke Wulf conducted dozens of drag studies on the FW-190G trying to reduce it's drag profile while still accomplishing the aircrafts mission.
Your wording very much betray's the fact you know the deception you are attempting to run.
All the best,
Crumpp
-
After some searching I found out that the tested plane was a Fw 190A-3 (Wnr. 511). It was tested in various configurations depending the purpose of the test, as an example below is one tested configuration (from FWB. Nr. 06006, there is several pics on other configurations too); profiled pylons and ETC in the centerline (with wheel doors, not an operational configuration). In the case of the report UM21 1, the drag at Cd0 is lower than in any configuration tested for the FWB.Nr. 06006 so it's certain that that the tested plane was in the clean configuration.
gripen
(http://personal.inet.fi/koti/temp0/pic2.jpg)
-
After some searching I found out that the tested plane was a Fw 190A-3 (Wnr. 511). It was tested in various configurations depending the purpose of the test, as an example below is one tested configuration (from FWB. Nr. 06006, there is several pics on other configurations too); profiled pylons and ETC in the centerline (with wheel doors, not an operational configuration). In the case of the report UM21 1, the drag at Cd0 is lower than in any configuration tested for the FWB.Nr. 06006 so it's certain that that the tested plane was in the clean configuration.
(http://www.onpoi.net/ah/pics/users/503_1112546305_pic2.jpg)
No Gripen,
Those are the Junkers Trager. Which were one of many different types tested. Only three types saw operational use.
(http://www.onpoi.net/ah/pics/users/503_1112537888_junkerstrager.jpg)
So beyond a shadow of a doubt you have found one of many drag studies on the FW-190G and are attempting to pass it off as a 190A.
From your demonstrated level of WWII aircraft design knowledge this can hardly be passed off as simple mistake.
All the best,
Crumpp
-
Dear Crumpp,
Why don't you stop your never ending personal attacks. I have given the reference, get the report and check if it's a drag study for the Fw 190G.
gripen
-
Parts of the translated version of Gripen's doc.
(http://butch2k4.free.fr/fwpolar1.jpg)
(http://butch2k4.free.fr/fwpolar3.jpg)
-
Thanks Butch!
Gripen the clean configuration of the FW-190G has a MUCH higher drag profile than the FW-190A fighter varient. This is due to intake changes and the absence of adjustable cooling gills. You can verify this from the level speed measurements.
The aircraft was designed to pull high loads for long distances.
All the best,
Crumpp
-
Well, those translations confirm pretty much all I said about the systematical error in the early measurements in the Chalais-Meudon.
The tested plane was a Fw 190A-3 as can be confirmed from the pictures so it had the adjustable cooling gills.
gripen
-
Well, those translations confirm pretty much all I said about the systematical error in the early measurements in the Chalais-Meudon.
Yes Gripen for the FW-190G for this report. Not for your implication that all FW-190 drag polars are wrong but this one.
It has no bearing on the FW-190A drag polars from both earlier and later test's. I have drag polars on FW-190A's from FW-190V5g to the FW-190A8/A9 (801TH). For the FW-190A series the polar agreement seems pretty good.
For your report on the FW-190G:
The prototype FW-190G was taken from a FW-190A3. The first production varients were the FW-190A4/U8, later blanket redesignated as FW-190G1.
No it does not have adjustable cooling gills and it does have a different intake.
Here is the part numbers for the tropical cowlings. Notice the G series is different. Why is that? It had a different intake!
(http://www.onpoi.net/ah/pics/users/503_1112574279_gintake.jpg)
The drag is totally different from an FW-190A series.
I am sure you already knew this though before you even posted.
Gripen says:
Why don't you stop your never ending personal attacks.
Coming from you this is a laughable comment. Maybe you should read your past post's starting with the most recent ones.
All the best,
Crumpp
-
Originally posted by Crumpp
Not for your implication that all FW-190 drag polars are wrong but this one.
I'm saying that the tests made in the Chalais-Meudon before the error was discovered seem to have a systematical error due to wrong correction factors. The FWB. Nr. 06006 data was tested in the July 1943 and the report is dated August and it has the mentioned error.
Originally posted by Crumpp
The prototype FW-190G was taken from a FW-190A3. The first production varients were the FW-190A4/U8, later blanket redesignated as FW-190G1.
The tested plane was a standard Fw 190A-3 and same airframe was used for wide range of tests. Please get the report.
Originally posted by Crumpp
Coming from you this is a laughable comment. Maybe you should read your past post's starting with the most recent ones.
When I bring in the data, you inform me that I can't read the plot, then you inform me that there is nothing wrong with the other plot despite the documentation says so, then you inform me that my translation is bad and so on...
When butch2k brings in the same data you thank him...
gripen
-
The tested plane was a standard Fw 190A-3 and same airframe was used for wide range of tests. Please get the report.
It's not the same plane, Gripen. It's the prototype FW-190G. Completely different aircraft with different characteristics. Documented and proven.
How many times does this have to be stated? It's in the report and it is proven by the photographs you posted.
Butch2K simply put what was in the report which is exactly what I said it was!!
Notice the wording of the document as well, MAY BE due to jet inclination. Facts are Focke Wulf never did figure out why they got such poor performance out of the FW-190G. The intake and lack of adjustable cooling fins are not the only changes. Quite a few changes were made to the engine as well including a completely different ignition harness. I suspect it was not one thing but a combination of minor performance robbing changes to set the aircraft up to haul large loads over long distances reliably.
The earlier test's from this report are effected!
Not every earlier drag polar as you tried to claim and got called out on!
In short it is useless as representative of the FW-190A.
All the best,
Crumpp
-
Originally posted by Crumpp
How many times does this have to be stated? It's in the report and it is proven by the photographs you posted.
This plane was tested in the wide range of configurations (including clean) and given the purpose of the tests (which can be figured out from the polar graph), it's certain that that the plane was in the clean configuration. Basicly it was a dedicated wind tunnel airframe which probably never flew.
Originally posted by Crumpp
Butch2K simply put what was in the report which is exactly what I said it was!!
Why don't you just ask from butch2k what kind test report it is.
Or better, get the report.
Originally posted by Crumpp
The earlier test's from this report are effected!
This is not what the report says, it says that earlier measurements from the another report (FWB. Nr. 06006) were effected.
"Vergleich mit früheren Messungen von Focke-Wulf"
It also tells how the error was discovered in the yet another measurements with the wing with symmetrical profile.
Besides such translation is unlogical, wrong and also gives an impression that the testers were idiots.
gripen
-
his plane was tested in the wide range of configurations (including clean) and given the purpose of the tests (which can be figured out from the polar graph), it's certain that that the plane was in the clean configuration. Basicly it was a dedicated wind tunnel airframe which probably never flew.
Gripen,
It was the FW-190G prototype.
Again proven by your pictures of it:
Gripens picture:
(http://www.onpoi.net/ah/pics/users/503_1112546305_pic2.jpg)
Junkers Trager for the FW-190G:
(http://www.onpoi.net/ah/pics/users/503_1112537888_junkerstrager.jpg)
Your contining to argue the obvious does not change this fact.
Even "clean", an FW-190G has much more drag than an FW-190A. The internal intake is different and the cooling fins are not adjustable.
Part numbers for ordering intake filters:
(http://www.onpoi.net/ah/pics/users/503_1112574279_gintake.jpg)
Again, your disputing the facts does not change them.
All the best,
Crumpp
-
Well, people believe the things they want to believe.
gripen
-
...what is the source of all this tabular data you get?
-Blogs
Originally posted by Badboy
The efficiency factor does work during that kind of maneuvering, what doesn’t work so well is the parabolic drag polar. The fact is that all the theory we have been discussing has evolved around the assumption that the relationships can be fitted with a parabolic polar. That turns out to be a very reasonable assumption for aircraft that have low AoA limits and depend exclusively on conventional lift. However, it is only approximately true, and in situations such as when other forms of lift are involved, or when the flow becomes turbulent near the edge of the envelope, it becomes more difficult to ignore. In those situations, it isn’t the value of e that needs to be modified, something other than a parabolic polar should be used.
The popularity of parabolic polars, and quadratic relationships and their appeal in aerodynamics stems from the fact that most of the resulting equations can be solved analytically, and generally they do produce good results, and errors are smoothed out to some extent within the coefficients. However, when you work with real aircraft, it isn’t always possible to use those simple relationships, and there are a number of alternatives. Firstly, you can model with non-quadratic relationships, but then the equations do not always have analytical solutions, (but they sometimes do) which doesn’t really matter because you can solve them using numerical methods, it just requires more computer time. There are other ways to match the real curves, very accurately, or at least with any degree of fidelity you wish, but it involves working with large quantities of tabular data. Both of those methods entirely resolve the issues of partial fit discrepancies for parabolic curves and enable very close predictions, even at the extremes of the envelope. I use those methods when ever I work with real aircraft, for EM analysis for example, but if you are analyzing what is happening in a simulation, such as AH for example, the traditional method is fine, because that is what is being used in the flight model anyway.
Hope that helps…
Badboy