Aces High Bulletin Board
General Forums => Aircraft and Vehicles => Topic started by: Stoney on February 28, 2008, 07:39:37 PM
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The "P-51...Awful" thread was starting to get cluttered, so I thought I'd start a new one for this.
I've got root/tip chord data for the Pony courtesy WW. Anyone have root/tip chord data for the Tempest and LA-7?
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Lets hope they also fix the 51's flaps and Fm with this next update.
Time will tell.
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I'm looking for the lengths in inches or cm or meters of the root and tip chords for the Tempest and LA-7.
That's it.
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can you scale off this?
(http://www.btinternet.com/~fulltilt/la7topbot2.GIF)
NACA root 23016
NACA tip 23010
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Hawker Tempest Root: H/1414/37.5 (14%) Tip:H/1410/37.5 (10%)
http://www.ae.uiuc.edu/m-selig/ads/aircraft.html
-C+
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Originally posted by Tilt
can you scale off this?
NACA root 23016
NACA tip 23010
Possibly. Just don't let anyone give me the Red A** when I only use a close approximation :)
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Originally posted by Charge
Hawker Tempest Root: H/1414/37.5 (14%) Tip:H/1410/37.5 (10%)
http://www.ae.uiuc.edu/m-selig/ads/aircraft.html
-C+
I've got all the airfoils. Just need root chord and tip chord.
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Originally posted by Tilt
can you scale off this?
From the drawing, I got an approximate of .84 meters at the outboard aileron edge (~tip chord) and 2.58 meters at the joint of the wing and wing fillet at the fuselage (~root chord). Anyone have any grief with these numbers?
EDIT: I did the measurements over again, as the best characteristic length for wing Reynolds Numbers is Mean Aerodynamic Chord. Using my CAD, I dimensioned the drawing and came up with a MAC of 2.26 meters, given a reference root of 2.7 meters, and a reference tip of .93 meters.
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Not until you bother to do the math, no.
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http://www.ajgs.com/Avaition-Protected/MACLength11ajgs.htm
or
http://www.nasascale.org/howtos/mac-calculator.htm
?
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Originally posted by Tilt
http://www.ajgs.com/Avaition-Protected/MACLength11ajgs.htm
or
http://www.nasascale.org/howtos/mac-calculator.htm
?
Don't know where they got their formulas, but Dr. Dan Raymer gives:
MAC = 2/3*Croot*((1+lambda+lambda)/(1+lambda))
Where:
MAC = Mean Aerodynamic Chord
Croot = Root Chord
Lambda = Taper ratio
Given that the taper ratio of the LA-7 planform is so high (~30%), it moves the MAC much closer to the root (and with a higher % of root chord) than on wings with more common taper ratios between 40-50%. Furthermore, continuing on (although not needed for my experiment), you use the following equation to determine the length in span, from the root, at which the MAC occurs:
Y = (b/6)*[(1+2*lambda)/(1+lambda)]
Where:
Y = Distance from wing centerline outboard
b = wingspan
[Taken from Aircraft Design: A Conceptual Approach]
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These are the LaGG-3 dimensions measured by the Germans. Should be pretty much the same as La-7, only minor differences in the wing root.
(http://www.onpoi.net/ah/pics/users/852_1204370254_lagg-3siipi.jpg)
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Neat gripen..........
La5FN and La7 did not have the outer fuel tanks (but the inners were slightly bigger) and the fuselage was wider on the Radial engined Lavochkins. Other wise the wing would be identical.
never occurred to me before but I guess the split in the flaps would have been across the wing joint.
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It's from a German study on wooden structure of the LaGG-3, there is also some pictures of early La-5s. It's a large report but mostly boring material analysis, however, there is some good info on details.
At least in pictures of Finnish LaGG-3s flaps go across the wing joint but it actually has a joint at the point outer wing starts. Below is the LaGG-3 series 35 "57" after capture, later it became LG-3 in FAF.
(http://www.onpoi.net/ah/pics/users/852_1204410986_lagg-3laippa.jpg)
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(http://www.btinternet.com/~fulltilt/la7frontgear.GIF)
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Flap splits seem pretty much in line with the joint.
btw
The front view on this sketch shows the engine exhaust vanes open either side of the fuselage. 2-3 o'clock and 9-10 o'clock.
When these were open on the La5FN (together with the front cooling vanes open to allow full engine cooling air) max speed was compromised by upto 45 to 50 km/hr and best turning time increased by upto 1.5/2 secs.
Despite the La7's better total aerodynamics (drag wise) a similar effect must of occurred.
In effect WEP was not just increasing revs it was also feathering the entire engine cooling system for minimum drag.
Lavochkin pilots truely had to fly with one eye always on the cylinder head temperature guage.
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Finished the Reynolds Numbers tonight. After this, I'll start compiling the lift/drag polars. There will be quite a few.
Reynolds Numbers were generated using the formula:
Rn=p*V*l/u
Where:
p = Density in slugs
V = Velocity in feet/sec
l = Characteristic length (used Mean Aerodynamic Chord)
u = viscosity in slugs
Method is to compare both the P-51D and LA-7 airfoils at 5 different speeds, at 5 different altitudes.
Speeds:
150 MPH
200 MPH
250 MPH
300 MPH
350 MPH
400 MPH
Altitudes:
Sea Level
5,000 feet
10,000 feet
15,000 feet
20,000 feet
I couldn't find the MAC for the Tempest, so I didn't include it. Given the number of polars I'm going to have to generate, I probably did myself a favor :)... Once I have all of the polars generated, I'll compare coefficients of lift, and drag, then compute lift/drag and plot that as well. These will not be exact, as I don't know how to accurately scale these airfoils to what they would be at the MAC for each aircraft, based on airfoil taper for both. So, what will be represented is the 23015 (15% thickness) for the LA-7 (which is probably close to the % thickness at the MAC), and the BL 17.5 (17.5% thickness) for the P-51D. The Pony airfoil will probably be 2-3% thicker than it is in real life, so I expect slightly higher lift and drag coefficients. Since I'm not espousing anything with this test, merely creating some data for comparison, the accuracy of the Pony airfoil should be close enough for us. The polars will be generated using XFoil96, using the Reynolds Numbers shown on the chart below, and their corresponding Mach numbers for the altitudes shown. Altitudes are assumed to be standard day. Roughness in XFoil will be set to 9.
The chart posted below shows each speed block, broken down by altitude, showing density, velocity, length, and viscosity for each condition. Two columns display the computed Reynolds numbers, at the far right, one each for the P-51D and LA-7.
(http://i125.photobucket.com/albums/p61/stonewall74/ReynoldsNumbersComparisons.jpg)
Once the polars are generated, I'll be back.
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1st Results:
(http://i125.photobucket.com/albums/p61/stonewall74/Cl_Comparison_SL150MPH.jpg)
(http://i125.photobucket.com/albums/p61/stonewall74/Cd_Comparison_SL150MPH.jpg)
(http://i125.photobucket.com/albums/p61/stonewall74/LIFT_DRAG_SL150.jpg)
For some reason, I was getting some faulty numbers for the P-51 above 15 degrees angle of attack. So, I left off everything above that number for both aircraft. Obviously, its easy to see the considerable advantage the P-51 airfoil possesses at cruise alphas. To reduce the amount of work for the comparison (this took me 3 hours), I'm going to pick an altitude for each speed. So, the 200 mph test will be at 5K, the 250 mph test at 10K, the 300 mph test at 15K, and the 350 mph test at 20K. I'll leave off the 400 mph test unless there's either some trend that justifies it, or on request.
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La7 Cy, Cx at various AoA at sea level at 40.5m/sec
(http://www.tilt.clara.net/vvscurves/PAGE1-18.JPG)
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Any one know what this is for???
(http://www.tilt.clara.net/vvscurves/PAGE1-19.JPG)
Pitching moment v AoA ?
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I could ask for a translation if no one else does it first..
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Where is fariz when you want him?
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Any one know what this is for???
Pitching moment v AoA ?
That's what it looks like to me. Its presented differently than I'm used to seeing it, say like the typical NACA graphs, but I'd say pitching moment.
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they need to make the p51 turn as it did in WW2. But if they did everybody would be flying them and i would get shot down more then i do allready :cry
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Stoney,
You have calculated profiledrag polars while Tilt's data shows the polar of the entire plane so these datasets are not comparable.
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Stoney,
You have calculated profiledrag polars while Tilt's data shows the polar of the entire plane so these datasets are not comparable.
Except to show how far the two can be apart?
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Tilt,
The profile drag coefficient is a two dimensional (2D) drag coefficient of a wing section while the the polar you posted shows three dimensional (3D) drag coefficient of the entire airframe. To give some picture how far these are from each other, you can read from your chart that at Cl 0 (Cy) the La-7 had Cd value (Cx) around 0,025 while the theoretical 2D drag coefficient of the NACA 23015 wing section (at MAC) is around 0,006. There are formulas to convert 2D data to 3D using wing geometry but even that does not make the data directly comparable because the effects of the fuselage and tail surfaces are still missing.
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Stoney,
You have calculated profiledrag polars while Tilt's data shows the polar of the entire plane so these datasets are not comparable.
At what point did I represent that I was doing anything other than "airfoil comparisons"? Tilt added some of those charts to flesh some things out, and that's ok.
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Stoney,
I'm not claiming that you have stated otherwise, sorry if you got such impression. However, these are more or less theoretical values; in practice in flight measured profile drag coefficients of the P-51 were far higher than wind tunnel measured. See the following NACA report on the XP-51:
http://hdl.handle.net/2060/19790073957 (http://hdl.handle.net/2060/19790073957)
To give some idea about the difference in profile drag coefficients at Cl 0,15:
Your data: about 0,0045
From the report:
Unfinished: 0,0070
Factory Finish: 0,0063
Factory Finish + sanding the insigna: 0,0059
Special smoothing by filling and sanding: 0,0053
So even in the best case with special finish, the profile drag coefficient was nearly 20% higher than theoretical wind tunnel value in the case of the P-51. Note that in the case of the NACA 23000 series profiles there probably was very similar difference; in the case of the Fw 190A-5 the Germans measured profile drag coefficient 0,0089 at Cl 0,2 while theoretical value was 0,0067 (in the case of the Bf 109B they measured 0,0101 vs 0,0068 at Cl 0,2).
Notable thing is that in the case of profile of the Mustang, so called "laminar flow bucket" disapeared with standard finish ie the polar shape was in practice very similar with older profiles.
In other words my opinion is that your comparison is more or less theoretical and valid only in the perfect conditions. Tilt's approach by using polars of the entire airframe is much better for comparing airframes.
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Tilt's approach by using polars of the entire airframe is much better for comparing airframes.
It was just some data I had . I have insufficient knowledge to form an "approach". But I see what you mean :) I think Stoney was going thru the theoretical approach to form a comparison rather than a model of reality. But then I really do not know what I am talking about here................. :rolleyes:
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Tilt,
It's just that using the polars of the entire airframe give quite bit different results than the wing section data. Below is the polars determined in the wind tunnel with a 1:3 scale model of the P-51. Note that I admit right away that scale models tend to give lower drag values than the real one in the full scale tunnel like the La-7 in the T-101 in your data so following is just an presentation how to use the polars:
(http://www.onpoi.net/ah/pics/users/852_1205169916_p51polar.jpg)
To get the normal form of the drag polar:
Cd = Cd0 + (CL^2/(pi*AR*e)
I quickly calculated the value of the e being 0,87 for the P-51 (using just three samples from the above polar) and from your data I got 0,84 for the La-7 (I used the better one of the two between Cl 0,2 and 1,0, I don't know what's the difference between these). So polars can be presented:
P-51 => Cd = 0,02 + (CL^2/(pi*5,8*0,87)
La-7 => Cd = 0,025 + (CL^2/(pi*5,5*0,84)
And graphically:
(http://www.onpoi.net/ah/pics/users/852_1205169941_lavs51polars.jpg)
And further as lift to drag ratio:
(http://www.onpoi.net/ah/pics/users/852_1205169963_lavs51clpercd.jpg)
So in practice there was no dramatic difference between these planes but these are just examples how to use the polars; my readings are just quick samples. In real life the Cd0 was around 0,022 for the P-51B in the full scale tunnel and value of the e was probably a bit lower (I tend to use numbers around 0,75 in my own calculations).
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Stoney,
I'm not claiming that you have stated otherwise, sorry if you got such impression. However, these are more or less theoretical values; in practice in flight measured profile drag coefficients of the P-51 were far higher than wind tunnel measured...Notable thing is that in the case of profile of the Mustang, so called "laminar flow bucket" disapeared with standard finish ie the polar shape was in practice very similar with older profiles...In other words my opinion is that your comparison is more or less theoretical and valid only in the perfect conditions. Tilt's approach by using polars of the entire airframe is much better for comparing airframes.
OK. To be honest, since we're simply comparing airfoils, I could have made the comparison to any of the NACA 23000 users out there, as it was the predominant airfoil used for the period. By using the Reynolds numbers that were associated with the respective geometries of the two aircraft, I thought it would be a little more realistic than simply a side by side comparison at theoretical reynolds numbers.
Since the drag coefficient for the aircraft includes more than simply the wing, obviously there are other factors involved. What can be said though is, comparitively, the Pony's airfoil is more efficient at its design lift coefficient than the NACA 23015, trim drag comparison not withstanding. (I was going to factor in pitching moments in the comparison, but didn't trust the XFoil data I got from the analysis).
But, like I said, I wasn't comparing the airframe--I tried to make that clear in the disclaimer posted early before the data.
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I wish I understood this................
However is there now enough data to dial in a comparison between the well documented P51 and the little documented La7 in particular with respect to acceleration and high speed dive characturistics.
btw re the two curves
The data is accompanied with photographs showing the ac at AoA's 2, 6, 8, 12, 16, 18, 20 degrees however there are 3 photographs per AoA
1 shows a standard La7 with engine cooling vanes fully feathered.
2 shows an La7 with additional booster bulges (on either side of the fuselage) again with engine cooling vanes fully feathered.
3 shows the same ac as 2 but now with cooling vanes open. (more drag)
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I wish I understood this................
However is there now enough data to dial in a comparison between the well documented P51 and the little documented La7 in particular with respect to acceleration and high speed dive characturistics.
What Gripen is talking about is the difference between a 2 dimensional airfoil comparison (what I'm showing), and the 3 dimensional wind tunnel tests. I'm looking at merely a slice of the wing at the mean aerodynamic chord. Obviously, since a wing is 3-D, there are other factors that the 2D comparison won't show. Gripen is correct in saying that an airfoil comparison doesn't take into account the 3D flow characteristics of an entire wing. Usually, a wind tunnel test will give a coefficient of lift/drag for the whole aircraft, while what I'm showing is merely that slice of the wing that exists at the MAC. That's why the maximum coefficients of lift are lower on the wind tunnel documentation.
The P-51D, as you alluded to, has plainly demonstrated documentation that shows the whole aircraft coefficient of drag. I cannot, through mathmatical terms, create that. Wind tunnel, flight testing, or computational fluid dynamics (CFD) software can.
What I was merely trying to do was show a comparison between the two airfoils only. I don't know if there's a published Cd for the LA-7. If there is, then we could plug them into the acceleration formulas for comparison. But, I can't create one--we'd have to find some documentation that shows one for a more precise comparison.
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However is there now enough data to dial in a comparison between the well documented P51 and the little documented La7 in particular with respect to acceleration and high speed dive characturistics.
Infact your data on the La-7 is very good for the airframe analysis; it's the real thing in the wind tunnel and the data shows large Cl range. It's tested just at one Mach and Reynolds number but it's still rather valid up to about Mach 0,4-0,5 because Cl/Cd ratio is quite constant up those speeds.
The data is accompanied with photographs showing the ac at AoA's 2, 6, 8, 12, 16, 18, 20 degrees however there are 3 photographs per AoA
1 shows a standard La7 with engine cooling vanes fully feathered.
2 shows an La7 with additional booster bulges (on either side of the fuselage) again with engine cooling vanes fully feathered.
3 shows the same ac as 2 but now with cooling vanes open. (more drag)
I don't really know if the vanes are the reason for the difference between the curves because the effect of the vanes should show up at all speeds as increased drag. However, the dragcurves are the same up to Cl (Cy) 0,4 or to the AoA of 4deg. My first impression was that the difference is caused by the automatic slots but that is just a quess. Perhaps some one who can Russian could translate graphs?
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Does anyone know if the AFDU trials or any turning performance trials during the war actually tested aircraft with their flaps down at varying degrees?
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However, these are more or less theoretical values; in practice in flight measured profile drag coefficients of the P-51 were far higher than wind tunnel measured.
Notable thing is that in the case of profile of the Mustang, so called "laminar flow bucket" disapeared with standard finish ie the polar shape was in practice very similar with older profiles.
Well, if I ratchet up the roughness used in the calcs, both airfoils show roughly the same comparison, albeit with much higher drag counts. The shape of the graphs are not quite as drastic, but maintain the same basic shape. I think it is safe to say that even though it doesn't match the theoretical values shown above, it still points to the P-51 profile being much more efficient on-design. The NACA 23XXX series were the most prevalent airfoil of the war, and certainly performed Yeoman's work for many aircraft of all sides in the war. Interestingly enough, as alluded to above, this doesn't take into account trim drag. The 23XXX was popular for many reasons, chief among them the lack of high pitching moments, while the P-51D airfoil had much higher pitching moments. Perhaps if a trim drag comparison was conducted, the gap between the two on-design would narrow appreciably.
<S>
Stoney
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Well, if I ratchet up the roughness used in the calcs, both airfoils show roughly the same comparison, albeit with much higher drag counts. The shape of the graphs are not quite as drastic, but maintain the same basic shape.
Below is one set of Naca 66 series section polars from "Theory of wing sections" by Abbott and von Doenhoff. I choosed this one because it just happened to be quite similar as profile used in the Mustang. Notable thing is that in the case of the smooth profile the air flow stays laminar at quite large part of the wing at low Cl values (say below 0,1). However when the Cl increases, the transition point from laminar to turbulent starts to move forward at certain point increasing drag as can be seen between Cl value 0,1-0,3 and at higher Cl values transition point stays again quite constant causing slower increase of the drag. This is the phenomena which causes so called "laminar bucket".
In the case of the Standard roughness, the airflow is allready turbulent at most of the wing even at low Cl values so there is no sudden movement of the transition point and no "laminar bucket" either. Therefore the basic shape of the polar should not be the same as in the case of the smooth surface.
(http://www.onpoi.net/ah/pics/users/852_1206627854_naca66.jpg)
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In the case of the Standard roughness, the airflow is allready turbulent at most of the wing even at low Cl values so there is no sudden movement of the transition point and no "laminar bucket" either. Therefore the basic shape of the polar should not be the same as in the case of the smooth surface.
Among most resources I've read, the NACA standard roughness figures are not "standard" with respect to service conditions of wings. Service condition is certainly worse than the theoretical values, but not nearly as rough as the standard roughness.
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Among most resources I've read, the NACA standard roughness figures are not "standard" with respect to service conditions of wings. Service condition is certainly worse than the theoretical values, but not nearly as rough as the standard roughness.
Well, we have measured data on the XP-51 (posted above), the normal condition profile drag values being near 50% higher than in the case of the smooth surface. There is also a lot of discussion on issue in the "Theory of Wing Sections".
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"Below is one set of Naca 66 series section polars from "Theory of wing sections" by Abbott and von Doenhoff. I choosed this one because it just happened to be quite similar as profile used in the Mustang. Notable thing is that in the case of the smooth profile the air flow stays laminar at quite large part of the wing at low Cl values (say below 0,1). However when the Cl increases, the transition point from laminar to turbulent starts to move forward at certain point increasing drag as can be seen between Cl value 0,1-0,3 and at higher Cl values transition point stays again quite constant causing slower increase of the drag. This is the phenomena which causes so called "laminar bucket"."
The effect of that is also well visible in Stoney's first set of plots with AoA to Cd and Cl as the effect of "the bucket" extends there too.
Gripen, could you post one of the 2300 series just for comparison. I have the book but not a working scanner...
-C+
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Here is the 23012.
(http://www.onpoi.net/ah/pics/users/852_1206707103_naca230123.jpg)
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Well, if you look at the standard roughness plots for almost every airfoil in the entire book, you'll see a Cdmin of something around .010, +/- about .2.
Given that, I have a hard time believing that the NACA standard roughness is representative of the "service condition". I agree that any airfoil in the service condition will fail to represent the theoretical values, but NACA standard roughness is anything but standard. Lednicer, Ribblet, etc. agree.
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Hm... I'm not refering on Standard roughness but on measurements on the XP-51 posted above. The service condition profile drag being about 50% higher than the ideal case so large amount of laminar flow is unlikely and therefore also appearance of the "laminar flow bucket" is unlikely as well. I posted the pages from Abbott and von Doenhoff just to show the phenomena, absolute values are irrelevant.
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Hm... I'm not refering on Standard roughness but on measurements on the XP-51 posted above. The service condition profile drag being about 50% higher than the ideal case so large amount of laminar flow is unlikely and therefore also appearance of the "laminar flow bucket" is unlikely as well. I posted the pages from Abbott and von Doenhoff just to show the phenomena, absolute values are irrelevant.
I thought that chart was from a wind tunnel model of the entire aircraft? Given that, I don't think you'd see a "bucket" that resembles the airfoil "bucket". I'll do a chart showing the effects of roughness on the P-51D root and we'll see how it plots.
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I'm refering this report (allready linked above):
http://hdl.handle.net/2060/19790073957
It's in flight measurements with the real XP-51 (2D profile drag). The polar (3D, entire airframe) comes from another tests on 1:3 scale model of the P-51:
http://hdl.handle.net/2060/19930092693
Generally the laminar flow bucket is really visible only in the 2D data.
In addition here is another report on profile drag:
http://hdl.handle.net/2060/19930092825
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Generally the laminar flow bucket is really visible only in the 2D data.
Well, that's my point. The 2D data merely shows the Cd of the airfoil. The 3D data shows the Cd of the entire aircraft. Given all the other components of drag, I wouldn't expect to see a "bucket" for the entire aircraft. The "bucket" should only be present for the airfoil.
Perhaps we're misunderstanding each other, as my only observation was that the P-51 airfoil was more efficient at design lift coefficients than the 23XXX series. The airfoil certainly is not the only part of the entire airframe to consider with respect to the Cd of the entire aircraft.
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The 2D data merely shows the Cd of the airfoil. The 3D data shows the Cd of the entire aircraft. Given all the other components of drag, I wouldn't expect to see a "bucket" for the entire aircraft. The "bucket" should only be present for the airfoil.
My point is that the measured far higher than ideal profile drag in the case of the real XP-51 indicates that there was no large amount of laminar flow in the service condition (see the first and the third linked reports). In other words there was no laminar flow bucket in 2D nor 3D. Basicly similar phenomena as seen in the case of the NACA 66 with standard roughness.
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the measured far higher than ideal profile drag in the case of the real XP-51 indicates that there was no large amount of laminar flow in the service condition (see the first and the third linked reports). In other words there was no laminar flow bucket in 2D nor 3D. Basicly similar phenomena as seen in the case of the NACA 66 with standard roughness.
Well, I looked through all 3 reports again, just in case I missed something. I noticed the table that listed the single Cd at differing finishes, but that was a single point Cd. Sure, the test values were higher, but a higher Cd at a single point does not mean that a bucket wouldn't have existed had an entire polar been generated in that roughness condition. The third report again tests the entire aircraft, and the Cd polar represents the entire airframe, and thus, there would be no bucket either.
Perhaps I'm just misunderstanding, but I'm not able to find in your references that a "bucket" didn't exist in the service condition.
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IMHO you don't seem to understand that they measured profile drag coefficient ie 2D in the first and third report:
http://hdl.handle.net/2060/19790073957 (this the report on in flight measurements of one section on the XP-51)
http://hdl.handle.net/2060/19930092825 (this is the report wind tunnel measurements of one section)
The results are directly comparable; the tested section was the same and the Re number in flight tests is close to wind tunnel tested. Given that in flight measured profile drag coefficient is far higher (about 50% higher in service condition), it's safe to assume that laminar flow did not happen in large degree in service condition. Therefore also it's safe to assume that there was no similar laminar flow bucket in the service condition as seen in the wind tunnel data.
Only the second report contains 3D measurements of entire airframe:
http://hdl.handle.net/2060/19930092693 (this is the report on wind tunnel tests on 1:3 scale model)
It's easy to see difference between 2D and 3D measurements because the absolute values of drag coefficient are very different: around 0,0045-0,0070 for 2D and around 0,02 for 3D at given Cl.
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Given that in flight measured profile drag coefficient is far higher (about 50% higher in service condition), it's safe to assume that laminar flow did not happen in large degree in service condition.
How do you make that assumption? In the first report, they only measured the Cd's at Cl's less than .4 which was the design lift coefficient for the airfoil. The plot for the "unfinished" profile begins to suggest a "bucket", even though they plotted it with a moving average. I suspect that if they had completed the plot through say, a Cl of 1.0, we might have enough information to determine whether or not a bucket existed in the service condition. The wind tunnel plots show a bucket on each and every chart.
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We know from the wind tunnel tests (third report) that when air flow stays laminar large part of the wing then the profile drag coefficient is around 0,0045 at Cl 0,15. However, in the flight tests measured profile drag (first report) is far higher in service condition, over 0,006 so we can conclude that the airflow is turbulent most of the wing. In other words there can't be large transition from laminar to turbulent when the Cl increases because airflow is allready turbulent most of the wing even at low Cl values. Therefore there can't be large laminar flow bucket either.
Another way to look the issue is to interpolate a polar using the third report assuming that most of the wing is allways turbulent (ie no large transition thus no large bucket). In that case the profile drag coefficient would had been a bit over 0,005 at Cl 0,15 which corresponds very well specially treated surface in the first report. There probably was some amount of transition even in the case of the unfinished wing, however, even in the case of the specially treated surface the transition was proabably far less than in the case of ideal smooth surface. Notable thing is that similar transition existed in some degree with with other profiles as well.