Aces High Bulletin Board
General Forums => Aircraft and Vehicles => Topic started by: Brooke on September 12, 2010, 03:15:30 PM
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Based on a discussion in another topic (which was not about P-51's).
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It started with this (which was originally about the P-38, but I foolishly put in one sentence about the P-51, which caused a derailment of the discussing into things totally about the P-51):
"Parasite" isn't the word I was looking for. The flow over surfaces is indeed turbulent -- the entire plane's surface was not designed to produce and could not produce laminar flow. No WWII airplane did that. Even laminar-flow wings on the likes of the P-51 probably didn't produce laminar flow once you factor in design variation, surface roughness, dirt, etc., and even if they did, they didn't produce laminar flow over the whole airframe. Also, just because there is a non-laminar layer of flow does not mean that huge increases in turbulence off a junction can't cause problems.
The tail buffetting that the wing fillet eliminated was not due to inadequate tail stiffening according to the following.
"Wind tunnel tests at Cal Tech established that tail flutter was the result of turbulent airflow created by the sharp juncture at wing and fuselage, and it was eliminated by a wing fillet that smoothed out the airflow over the tail." From American Aviation, by Christy and Cook, and you can see the excerpt here:
http://books.google.com/books?id=E6yzMq7Z-yIC&pg=PA178&lpg=PA178&dq=p-38+wing+fillet&source=bl&ots=lLWgRD_uDN&sig=yYV4Ff__D6DJeLVdQinv-gLdSy0&hl=en&ei=gNiLTMLFHo6msQOzuYCIBA&sa=X&oi=book_result&ct=result&resnum=7&ved=0CCUQ6AEwBg#v=onepage&q=p-38%20wing%20fillet&f=false
Other references saying the same:
The Lockheed P-38 Lightning, by Bodie. The marvelous and definitive book on all aspects of the P-38.
http://www.amazon.com/Lockheed-P-38-Lightning-Warren-Bodie/dp/0962935956/ref=sr_1_1?ie=UTF8&s=books&qid=1284234200&sr=8-1
http://www.fighter-planes.com/info/p38_lightning.htm
http://www.aviation-history.com/lockheed/p38.html
http://en.wikipedia.org/wiki/P-38_Lightning
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Oh... by the way.
The P-51 was the first laminar flow design ever implemented. You are correct in saying that it did not experience laminar flow... but only if you mean over the entire wing. Yes the P-51 was designed for laminar flow and yes it did experience laminar flow over 40-50% of its airfoil by design. Yes there were problems with reduced performance over theoretical and design potential... but to say that the P-51 never experienced laminar flow over any portion of its airfoil is not an honest statement. If the Mustang never experienced any form of laminar flow then it would never have the performance it did and does.
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but only if you mean over the entire wing.
Yes, that's what I mean
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In that case it doesnt matter. In having laminar flow over 40-50% of the surface area you are eliminating that much effective drag (turbulent) and because the layers are not separated until the point of minimum pressure the effective profile drag is much less whereas on an aircraft with turbulent drag over the entire wing surface (99.9%) the profile drag can in effect be as much as doubled. Also... the laminar airfoil of the P-51 eliminates the increase of the coefficient of drag for the majority of the effective AOA. What that means is that as a pilot increases the AOA the coefficient of lift increases but the coefficient of drag remains constant for 65% (approximately) of the effective AOA. This is why a P-51 can raise its nose as much as eight degrees (or a little more) and the drag remains the same as if the wing were level and the P-51 can use its E to zoom to great advantage. No P-51 should ever zoom straight up unless the pilot is willing to give up a great deal of his E to do it (the coefficient of drag will climb at a greater rate than on a Spitfire for instance).
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I would be willing to bet a large sum of money that a P-51 with wing at 8 degrees angle of attack has substantially more drag than a P-51 with wing at zero degrees angle of attack.
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Also, the following is directly to the point of what I first wrote: "Even laminar-flow wings on the likes of the P-51 probably didn't produce laminar flow once you factor in design variation, surface roughness, dirt, etc."
From http://yarchive.net/mil/laminar_flow.html
From: Charles.K.Scott@dartmouth.edu (Charles K. Scott)
Newsgroups: rec.aviation.military
Subject: Mustang, was it's wing really laminar flow? Very long
Date: 5 Feb 1997 14:13:04 GMT
Was the Mustang's laminar flow wing laminar or not?
That is a question asked often in several groups and recently, after
finishing "Pursue and Destroy" by Leonard "Kit" Carson, I believe I
have found a definitive answer.
Mr. Carson's credentials are that 1. He flew the Mustang in combat. 2.
He was an engineer who understood aerodynamics. 3. He was a test pilot
for a while after W.W.II. He goes into extreme technical detail while
telling about the Mustang and his career flying it.
Carson begins his analysis of the Mustang and it's laminar flow wing
back in the late 20's and early 30's when NACA, the National Advisory
Committee for Aerodynamics began it's research on airfoils, airflow,
and other aspects of flying. The airplane companies were in no
position to do this research because they did not have the money to
develop and build wind tunnels. He described airfoil research prior to
NACA as piecemeal, with many airfoils being developed by the OTLAR
method (Oh That Looks About Right, my words, not his)
It was during the thirties that NACA established the relationship
between turbulent flow and drag. Their measurements indicated that the
3/32 inch rivets heads and lap joints on the typical metal airliner
"dissipated" 182 horsepower. On one airplane they were measuring, they
found that a coat of paint cost the airplane 91 horsepower over the
same airplane with bare aluminum. They learned that mere dust, fine
sand or a piece of scotch tape "would cause the smooth laminar layer
next to the wing surface to jump over To a turbulent, high drag
condition."
Then, in 1938, in a wind tunnel designed to smooth out the airflow
through the tunnel (designed by Jacobs and Dryden, prior to this wind
tunnel, flow through the tunnel was too turbulent to test laminar
theories) a new type of airfoil was tested that set new and incredible
drag coefficients compared to any airfoil previously tested. It
recorded a drag coefficient of .003 "which was about half of the lowest
ever recorded before for an airfoil of similar thickness."
Further tests conducted in England "demonstrated that laminar flow and
a reduction of drag could be obtained for a considerable distance over
a smooth full scale wing."
This was in the wind tunnel, however, and it turned out that an
enormous gulf existed between test aircraft and the wind tunnel and
combat aircraft.
The following reasons were given by Carson explaining why in real life
laminar flow simply did not occur on the P-51's wing.
1. The effects of propeller Slipstream. Airflow within the arc of the
prop is very turbulent, "the whole fuselage and inboard section of the
wing next to the fuselage operate in that turbulent stream. Tests in
the Langley wind tunnel revealed that airflow within the arc of the
prop (the prop was 11 feet in diameter which meant that turbulent air
was encountered all the way out to within 13 inches of the inner gun
position) was "90 to 95 percent turbulent" (in other words non laminar)
2. Vibration: "Engine and propeller vibrations transmitted through the
structure will induce transition to turbulence." Tests indicated that
laminar flow on twin engine aircraft was greater with one engine
feathered than with both running. Engineers surmised that the lack of
engine/prop vibration on the dead engine side promoted laminar flow.
Honest, that's what the book said. Of course with both props turning,
more of the wing would be bathed in the prop slipstream which as has
been mentioned above, trips laminar flow to turbulent.
3. Airfoil surface condition: "Mud, dirt, ice and frost will induce the
transition to turbulent conditions." "Fuel truck hoses, ammo belts,
tools, guns and large feet in GI. shoes found the way to the tops of
wings" the scrapes and dents this servicing caused had negative effects
on laminar flow.
4. Manufacturing tolerances: "The Mustang was the smoothest airplane
around in 1940, but there is a practical limit in construction. We're
talking about surface roughness or waviness of .01 inches which will
cause transition to turbulence." (remember the afore mentioned dust
and scotch tape which was observed to trip airflow to turbulent). Some
aerodynamicists have stated that true laminar flow did not occur
outside the wind tunnel until the advent of Burt Rutan's Vary E-Z in
the early 70's with it's incredibly smooth fiberglass over carved foam
wing and aft mounted engine which of course kept the wing ahead of the
prop slipstream.
5. Wing Surface Distortion in Flight: Flight brings flight loads which
can and did distort the wing and cause ripples in the wing surface
which were fully capable of tripping the laminar flow to turbulent.
Carson went on to state: "The Mustang wing was a high lift
configuration, as well as low drag. . . the Mustang in squadron service
was not laminar to the same extent as the wind tunnel development
models. Not one day in the past 34 years (the book was written in 74)
has it performed in that manner for any or all of the reasons just
given."
So if it wasn't the laminar flow wing that gave it it's high speed and
extensive range, what was it?
The most prominent speed secret was the dramatic reduction of cooling
drag. Placing the airscoop on the belly just in front of the rear edge
of the wing removed it as far as was practicable from the turbulence of
the prop and placed it in a high pressure zone which augmented air
inflow. Tests in the wind tunnel with the initial flush mounted scoop
were disappointing. There was so much turbulence that cooling was
inadequate and some doubted that the belly scoop would work. The
breakthrough was to space the scoop away from the surface of the belly
out of the turbulent boundary layer of the fuselage. Further testing
showed that spacing it further out would increase cooling but at a cost
to overall drag. Various wind tunnel tests established the spacing at
the current distance which represents the best compromise between
spacing out from the turbulent flow of the fuselage, drag and airflow.
With the flow into the scoop now smooth and relatively nonturbulent,
the duct leading to the radiator/oil cooler/intercooler was carefully
shaped to slow the air down (the duct shape moves from narrow to wide,
in other words a plenum chamber) enough from the high external speeds
to speeds through the heat exchangers that allowed the flow to extract
maximum heat from the coolant. As the air passed through the radiators
and became heated, it expanded. The duct shape aft of the radiator
forced this heated and expanded air into a narrow passage which gave it
considerable thrust as it exited the exhaust port. The exhaust port
incorporated a movable hinged door that opened automatically depending
on engine temperature to augment the airflow. The thrust realised from
this "jet" of heated air was first postulated by a British
aerodynamicist in 1935. The realization of thrust from suitably
shaped air coolant passages is named after him and called the "Meredith
Effect". Some have said that at certain altitudes and at a particular
power setting the Meredith effect was strong enough to actually
overcome all cooling drag; this is not regarded as being accurate by
most aerodynamicists. It greatly contributed to overall efficiency of
the cooling system but never equaled or overcame cooling drag.
Combine the low overall drag of the Mustang with significantly greater
internal fuel tankage than either the Spitfire, Messerschmitt or
Focke-Wulf 190 and you can easily see how it could fly so far. Add the
two 105 gal external wing tanks and the Mustang was fully capable of
flying to any target the heavy bombers could attack in the ETO. Kit
Carson mentioned that he flew more than 35 missions during which he was
in the cockpit for more than 5 hours.
Finally, Carson was interested to find, while reading flight test
reports in research for his book, that the quoted top speed for the
P-51B was less than what was attained during test flying. The
information is as follows:
Report: NA-5798
Title: "Flight Test Performance for the P-51B-1
Date: January, 1944
Test Weight: 8,460 lbs
High Speed: 453 mph true airspeed at 28,800 feet at 67" HG and 1298 HP,
war emergency power, high blower, critical altitude.
The quoted top speed for the B model Mustang is 440 mph.
I can only speculate that it is likely the test airplane used in the
above mentioned flight was a well maintained and unblemished Mustang.
It's probable that the actual combat aircraft would not be able to
quite equal that performance. Never the less, Carson notes this
information and concludes with the following:
"It's easy to see why many pilots preferred the P-51B, including
myself, even if it did have only 4 guns and the "birdcage" canopy. If
you can't hit'em with 4 guns, two more aren't' going to make your aim
any better."
Corky Scott
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I would be willing to bet a large sum of money that a P-51 with wing at 8 degrees angle of attack has substantially more drag than a P-51 with wing at zero degrees angle of attack.
Your going to lose that pile of money... but your large sum probably equates to about $15.
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Great post Brooke. :aok
Really makes you appreciate the fact that some ground crews went the extra mile to clean and polish planes to help get they're favorite pilot back home.
:salute
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I would be willing to bet a large sum of money that a P-51 with wing at 8 degrees angle of attack has substantially more drag than a P-51 with wing at zero degrees angle of attack.
Just happen to have this handy from a comparison I did a few years ago...
(http://i125.photobucket.com/albums/p61/stonewall74/Cd_Comparison_SL150MPH.jpg)
This is for 150mph TAS at sea level. I couldn't run anything approaching max speeds because XFOIL doesn't like laminar airfoils and high R numbers, but the trend would be the same. For comparisons, the La-7 line was plotted using the 23015 airfoil used by just about every other non-British WWII fighter.
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As a footnote, I do not believe that the aerodynamic efficiencies of the P-51 were merely due to low cooling drag. While theoretical levels of laminar flow would never have been achieved in service aircraft, the relative comparison between the P-51 and the turbulent airfoils of peer aircraft, also in the service condition, would be consistent with the relative comparison of theoretical drag. Further, the P-51 was, from prop spinner to tail tip, optimized aerodynamically. The design was brilliant, in my opinion, in that regard. The laminar flow airfoil was just a piece of that overall aerodynamic optimization, but a significant one none-the-less.
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Another reason for the wing design was to increase the area capacity for guns, gear, and tanks in the center of the wing rather than near the front edge. The laminar flow design keep the wing’s internal space efficient, with sufficient lift, and minimal drag.
The word "laminar" definitely made the plane sound new and cool. Not far from advertising things like duel exhaust, super-hydromatic twin turbocharged multi staged intercooled intake/exhaust. Another popular feature that cracks me up is things like hood scoops, spoilers, auto-extending spoilers. Those devices may help in some cases, but for everyday driving they may make an automobile heavier with more drag, thereby, making the car slower. :rofl
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Oh boy. You know once I start answering this it is going to create a wall of text that you will neither read nor accept... but it has to be done.
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Which aircraft in WWII had laminar flow wings? P-51 I know and I seem to recall reading that the Tempest and Ki-84 did as well.
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This is why anecdotes are ignored.
First off you have to understand that the engineers at N.A. used all of the available data they could get their hands on and then they refined that information by making use of the wind tunnel at UCLA (actually I forget if it was UCLA or CALTECH... one of those CA schools). Most of the Universities of the day (like today) were teaching data that lagged behind laboratory discovery and so the wind tunnel was a requirement. A NA engineer remained at the wind tunnel day and night to make use of every available free moment that he could.
The stated reason number one against laminar flow is propeller slipstream. However anyone that has studied aerodynamics and propellor forces knows that there is a phenomena called 'slipstream contraction' that occurs aft of the propeller backplate. This was understood by NA engineers as it should have been by Mr Carson (the engineer that understood aerodynamics). The slipstream contraction effect is further influenced by the fuselage of the aircraft (P-51) to create a central vortex in circulation about the fuselage. You can see the engineered corrections to enhance the contraction in the design of the wing inward of the gun positions. There was also a geometric change in the airfoil from center to wingtip implemented to avoid the need for physical washout and to correct for the change in chord length. This reduced the effect of propeller slipstream to within 3-4 feet of the fuselage which you will see in the end was accounted for anyway. By the time of the P-51H this was no longer deemed necessary because of other changes to the shape of the fuselage and filleted areas (enhancing contraction further).
Number two reason is vibration which although true is more effective at created turbulent stream if it is the air that is vibrating. The effect of vibrating wind tunnels has led to trememdous efforts to reduce turbulence within the tunnel itself. A Mustang that effectively destroys its enemy through explosion will experience a loss of laminar flow. Further rebuttal of this is not possible because of the lack of details in the account. What twin engine airplane had a laminar wing? Most certainly the vast majority of aircraft of the era had laminar flow over only 2-3mm of the wings leading edge. I believe the test is altogether untrue due to the impracticality of testing pressures in real world flight... how did they know they EVER had laminar flow. Too many details are missing.
Point number three involves the wing surface. It is true that a film surface of .002" will restore laminar flow. It is also true that a single paint chip can destroy laminar flow over that portion of the wing but it does not destroy flow over the entire wing. This was discovered back in the days of the Mustang I. This is one reason you dont often see men climbing over the leading edge of a P-51. The wing root and the trailing edge of the wing behind the ammunition bays have zero effect upon laminar flow. This is also one reason most Mustangs after the B/C era are unpainted. Remember the only concern with the Mustang is the first 40-50% of the airfoil.
Point number four is fairly accurate. True laminar flow (over the majority of the wing) did not come about until modern times. That does not mean that laminar flow exists only in modern times. Even the best wings today only achieve about 96% laminar flow. Hearing that confuses people into thinking that the limited laminars of the past (40-50% laminars) never achieved thier designed goals which is not true.
Point number five is probably accurate too however I dont believe I have ever seen a photo of a Mustang with wrinkled skin. Perhaps in heavy manuevers when the pilot is stressing not only his skin but the airplanes too and while at a very high speed indeed... but no.
Then the point moves toward the low drag radiator. I agree it did help tremendously.
HOWEVER he goes a little too far in comparing the fuel load of the Mustang with Spits and 109s. It seems he also forgot a few things about the law of physics concerning lift and weight vectors. Yes the Mustang carried more fuel and it also weighs more. However... please alert CorkyJr about those wing tanks I want some!
Then he slips up and forgets that the tests conducted on the 'B' model were with removed guns and armor.
As to the cl/cd graph... this one compares the P-51 versus the Spitfire (I believe 14 model). You can clearly see that the lift created by the Spitfire wing exceeds the lift of the P-51 but only after a given AOA is exceeded does the drag begin to effect the Mustang.
(http://i447.photobucket.com/albums/qq197/Chalenge08/P-51_6series_vs_Spit14_4series.png)
Now if you need graphics of a 40-50% flow versus a 96% flow in order to understand my points just ask.
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Which aircraft in WWII had laminar flow wings? P-51 I know and I seem to recall reading that the Tempest and Ki-84 did as well.
Just the Pony... According to Dave Lednicer, the Ki-84 shared the 23000 series airfoil with most of the other WWII fighters. The Tempest used a proprietary airfoil developed by Hawker, but if the plots I've seen of it are correct, it isn't going to be a laminar airfoil. The P-63 used one, but I don't really count it in the mix.
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As to the cl/cd graph... this one compares the P-51 versus the Spitfire (I believe 14 model). You can clearly see that the lift created by the Spitfire wing exceeds the lift of the P-51 but only after a given AOA is exceeded does the drag begin to effect the Mustang.
Do you have the data points for those drag polars? That chart looks like an illustrative comparison between a polar that demonstrates the "laminar bucket" and a turbulent airfoil. Its good for showing the concept, but can't be used for quantitative comparison. If you look at my graph, you can see the laminar bucket on the P-51 plot. However, every airfoil I've ever seen, except for those designed with very high design lift coefficients will have more profile drag at 8 degrees of alpha versus 0 degrees of alpha.
However, remember that effective AoA is computed from the relative wind. In a near 1G climb at high speed, it is possible to have a situation where there is 8 degrees of nose-up pitch on the aircraft, but the effective AoA is close to cruise AoA.
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Your going to lose that pile of money... but your large sum probably equates to about $15.
So the bet is whether this statement is true or not: "The P-51 at 8 degrees angle of attack has more drag than the P-51 at zero degrees angle of attack," where we are talking about real-world P-51's as in service in WWII. We'll give our wagered money to a mutually trusted third party and equally chip in to hire three aerodynamics experts to opine on it. Winner is paid by the third party based on decision of the panel.
How much do you want to bet?
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So the bet is whether this statement is true or not: "The P-51 at 8 degrees angle of attack has more drag than the P-51 at zero degrees angle of attack," where we are talking about real-world P-51's as in service in WWII. We'll give our wagered money to a mutually trusted third party and equally chip in to hire three aerodynamics experts to opine on it. Winner is paid by the third party based on decision of the panel.
How much do you want to bet?
I thought I had already nipped that one in the bud... :)
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I thought I had already nipped that one in the bud... :)
Only if Chalenge believes your modelling or opinion on it. I'll take bets on that one, too. :)
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Which aircraft in WWII had laminar flow wings?
The B-24 had the Davis Wing, which provided some amount of laminar flow.
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The B-24 had the Davis Wing, which provided some amount of laminar flow.
Its imprecise to characterize "wings" as laminar flow. Laminar flow is not a characteristic of wing design, it is a characteristic of airfoil design. The Davis Wing was unique because it emphasized aspect ratio, chord thickness, and reduced wing area, not because it generated laminar flow. The airfoil used on the Davis Wing created the laminar flow associated with the wing. In most circumstances though, the amount of laminar flow created by the airfoil would still put it in the "turbulent" category.
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Only if Chalenge believes your modelling or opinion on it. I'll take bets on that one, too. :)
Well, an actually, you don't even need the airfoil analysis I did for that, as you can derive the induced drag coefficient of the P-51 at 8 degrees AoA without it. If the Cdi is greater at 8 AoA, then it must create more drag at 8 degrees AoA than at zero. Of course, you know this Brooke. (Actually, the design lift coefficient of the P-51 gives it zero lift at ~2 degrees AoA)
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Stoney I believe the problem you are having is that you are unable (due to limitations of XFoil) to get away from the region of flight in which vortex drag (the poorly named "induced drag") is the overwhelming drag on aircraft. What you cant see in the comparison chart I submitted is the AOA points and airspeeds (constant CV) which were in the range where profile drag and parasite drag overwhelm vortex drag significantly.
Mapping a laminar profile at such a low speed (outside of radio control airfoils) is a little unfair and not at all practical considering the full range of these aircraft.
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[edited to get rid of debating terminology]
The Davis Wing was unique because it emphasized aspect ratio, chord thickness, and reduced wing area, not because it generated laminar flow.
The Davis Wing emphasized less drag for a given amount of lift. To do that, it had those characteristics including a laminar-flow design -- intentional or not, I have no idea, but without the laminar-flow characteristic, according to the following, the Davis Wing probably would not have been adopted, so that part was perhaps most important.
http://en.wikipedia.org/wiki/Davis_wing
In most circumstances though, the amount of laminar flow created by the airfoil would still put it in the "turbulent" category.
Yes -- that was my point above. Dirt, paint, bugs, footprints, design variation, skin wrinkles, scratches, airframe vibration, turbulence from the propeller, etc. all mean that in practice (not theoretically, or with a perfectly smooth version in a perfectly clean lab) those laminar-flow wings (or "wings based on laminar-flow airfoils" or whatever folks want to call them) did not fully produce laminar flow. However, they did seem to have significantly lower drag than other non-laminar-flow designs and so were good wings in that regard. (That might be because they still had laminar flow for at least a portion of their chord, as Chalenge says -- that is not the point I was debating with him.)
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Pretty nice definition of laminar flow wing here too: http://www.aviation-history.com/theory/lam-flow.htm
-C+
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Brooke I will agree with you that the majority of wings instill additional profile drag due to flow seperation and increased turbulence as the AOA increases. This is also true of the P-51s laminar wing but only after a point. I believe I recall that point being 8.8 degrees with the coefficient of lift dropping below a Spitfire at 10.2 degrees and drag increasing dramatically thereafter. The reason most people cannot picture this as possible is due to the image of general flow coming from directly ahead and level with a zero AOA line which is not the case. In order to picture the effects of a wing upon open air flow (non-wind tunnel) you can think of a wing as being a rotating vortex like a tornado as seen from above. As air flows into the vortex it is curved by the approach of the varying pressure zones. This is not what actually happens with a wing (a rotating vortex I mean) but the effect of oncoming air flow is the same. For the majority of wings the stagnation point where approaching flow meets wing leading edge will drop below the leading edges center point significantly further than it does with a laminar design. I think if you see that much clearly you can see the reason that profile drag does not increase for a laminar airfoil until the low drag limitation is exceeded.
I also do not believe the B-24 had any laminar flow (beyond 2-3mm). Please site anything other than wikipedia if you have it. From what I recall of airfoil theory the Davis wing was designed with a smoothly increasing profile in such a way that even though turbulence occured it did not cause rapid separation and the onset of the separation bubble occurs well aft of the minimum pressure point. The problem with this design is the profile is so large it causes issues at high altitudes and probably the only fix would have been more horsepower. I belive this fact is born out in historical accounts as well as the performance in AH.
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[edited to get rid of debating terminology]
The Davis Wing emphasized less drag for a given amount of lift. To do that, it had those characteristics including a laminar-flow design -- intentional or not, I have no idea, but without the laminar-flow characteristic, according to the following, the Davis Wing probably would not have been adopted, so that part was perhaps most important.
http://en.wikipedia.org/wiki/Davis_wing
Yes -- that was my point above. Dirt, paint, bugs, footprints, design variation, skin wrinkles, scratches, airframe vibration, turbulence from the propeller, etc. all mean that in practice (not theoretically, or with a perfectly smooth version in a perfectly clean lab) those laminar-flow wings (or "wings based on laminar-flow airfoils" or whatever folks want to call them) did not fully produce laminar flow. However, they did seem to have significantly lower drag than other non-laminar-flow designs and so were good wings in that regard. (That might be because they still had laminar flow for at least a portion of their chord, as Chalenge says -- that is not the point I was debating with him.)
I'm not trying to be obstinate Brooke--its important to understand the difference between the characteristics of wing design and the airfoils that are used on them. Wings by themselves shouldn't be considered as "laminar" as wing design does not generally contribute towards laminar or turbulent flow.
Wing design considers planform shape, aspect and taper ratios, wing tip design, wing to fuselage placement, anhedral or dihedral, sweep, etc. The P-51 wing design was, like the Davis wing, a remarkable step forward in aerodynamic development. They further increased the efficiency of both wings by using airfoils with low profile drag.
Airfoil use introduces the concepts of "laminar" versus "turbulent" airfoils. All airfoils experience laminar flow. Generally speaking, any airfoil that can maintain theoretical laminar flow beyond the 30% chord region is considered a "laminar" airfoil. Any airfoil who's boundary layer is interrupted short of that 30% region is considered a "turbulent" airfoil.
The P-51 has a laminar airfoil, regardless of whether or not it achieves it in the service condition, because it fits the definition. That same airfoil on a wing made of composites would definitely generate laminar flow, even in the service condition. You are correct that service P-51s suffered from dings and dents that would have interrupted laminar flow in flight, but it still is considered, from a design perspective, to have a laminar airfoil.
The B-24, on the other hand, achieved most of its low drag characteristics as a result of wing design. I haven't analyzed the airfoil in XFOIL, but from the Wiki description you linked, it would appear that the airfoil itself was a very efficient, turbulent (by definition), airfoil.
This may sound like semantics, but its not.
@Chalenge, I'll respond to your post later today...
(good discussion regardless...)
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I came by a MIT lecture video where Dale D. Myers, who worked on the P-51 design, mentions that scoop solution before starting his actual lecture about space shuttle systems engineering.
Here's the link if you're interested: http://www.youtube.com/watch?v=iiYhQtGpRhc (http://www.youtube.com/watch?v=iiYhQtGpRhc)
He talks about the Mustang briefly at about 16:00 mins into the video.
So if it wasn't the laminar flow wing that gave it it's high speed and
extensive range, what was it?
The most prominent speed secret was the dramatic reduction of cooling
drag. Placing the airscoop on the belly just in front of the rear edge
of the wing removed it as far as was practicable from the turbulence of
the prop and placed it in a high pressure zone which augmented air
inflow. Tests in the wind tunnel with the initial flush mounted scoop
were disappointing. There was so much turbulence that cooling was
inadequate and some doubted that the belly scoop would work. The
breakthrough was to space the scoop away from the surface of the belly
out of the turbulent boundary layer of the fuselage. Further testing
showed that spacing it further out would increase cooling but at a cost
to overall drag. Various wind tunnel tests established the spacing at
the current distance which represents the best compromise between
spacing out from the turbulent flow of the fuselage, drag and airflow.
With the flow into the scoop now smooth and relatively nonturbulent,
the duct leading to the radiator/oil cooler/intercooler was carefully
shaped to slow the air down (the duct shape moves from narrow to wide,
in other words a plenum chamber) enough from the high external speeds
to speeds through the heat exchangers that allowed the flow to extract
maximum heat from the coolant. As the air passed through the radiators
and became heated, it expanded. The duct shape aft of the radiator
forced this heated and expanded air into a narrow passage which gave it
considerable thrust as it exited the exhaust port. The exhaust port
incorporated a movable hinged door that opened automatically depending
on engine temperature to augment the airflow. The thrust realised from
this "jet" of heated air was first postulated by a British
aerodynamicist in 1935. The realization of thrust from suitably
shaped air coolant passages is named after him and called the "Meredith
Effect". Some have said that at certain altitudes and at a particular
power setting the Meredith effect was strong enough to actually
overcome all cooling drag; this is not regarded as being accurate by
most aerodynamicists. It greatly contributed to overall efficiency of
the cooling system but never equaled or overcame cooling drag.
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The radiator scoop on the P-51 is what is commonly called a boundary layer scoop, the same design was used years later as a hood scoop on race cars. It creates less drag than many scoops for the same amount of frontal area and air intake, creating something of a ram effect. Ramming air into the radiator with an efficient low drag scoop, then using the heated air to create thrust is almost like free HP. The gain is from taking a necessary function that is usually pure drag, and lowering the drag, while increasing the thrust, so instead of being almost pure drag, it actually contributes thrust, with lower drag as well.
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I would be willing to bet a large sum of money that a P-51 with wing at 8 degrees angle of attack has substantially more drag than a P-51 with wing at zero degrees angle of attack.
Here's a chart. 8 degrees AOA corresponds to a lift co of about .75 with flaps up.
(http://beta.hitechcreations.com/pyro/p51ddrag.jpg)
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Stoney I believe the problem you are having is that you are unable (due to limitations of XFoil) to get away from the region of flight in which vortex drag (the poorly named "induced drag") is the overwhelming drag on aircraft. What you cant see in the comparison chart I submitted is the AOA points and airspeeds (constant CV) which were in the range where profile drag and parasite drag overwhelm vortex drag significantly.
Mapping a laminar profile at such a low speed (outside of radio control airfoils) is a little unfair and not at all practical considering the full range of these aircraft.
Well, the drag shown in most drag polars is profile drag, not induced drag. I shouldn't have said the drag polar proves Brooke's contention. But, given the formula:
Cdi = Cl^2 X pi X e X AR
Where:
Cdi = induced drag coefficient
Cl = lift coefficient
pi = 3.14...
e = Oswald's efficiency number
AR = Aspect ratio
Its obvious that any increase in lift coefficient (aka AoA) will result in a higher Cdi, and thus, induced drag must increase with effective AoA.
In 2-D modeling such as XFOIL, there is no induced drag as the 3-D effects of lift are not considered as the wing is considered to be of infinite span. What it shows is purely the drag created by increasing the AoA of the airfoil section. In order to create induced drag modeling, the entire wing must be considered. So, for pure comparisons between airfoils, the drag polars should suffice. As to the Reynolds number used, the drag numbers will only be lower, and would compare approximately the same--perhaps with the laminar airfoil showing even more of a low-AoA advantage, as they're performance generally improves with higher R-numbers. At that speed and using the MAC of the P-51, we're still talking about a mid-7 figure R-number (around 6,000,000 or so), which is not nearly as low as the mid-6 figure R-numbers (around 400,000 or so) normally associated with drag polars of RC airfoils.
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Just the Pony... According to Dave Lednicer, the Ki-84 shared the 23000 series airfoil with most of the other WWII fighters. The Tempest used a proprietary airfoil developed by Hawker, but if the plots I've seen of it are correct, it isn't going to be a laminar airfoil. The P-63 used one, but I don't really count it in the mix.
Stoney,
I't doesn't have much relevance to the discussion I know... but just for clarification,
I don't know for all airplanes, but I know that the 109 Fs and later did not have the same 2300 series airfoil that the 109Es had.
109E arifoil - NACA 2314(root)- NACA 2310(tip)
109F+ airfoil - NACA 2R1 14.2(root) - NACA 2R1 11.35(tip)
how close is a NACA 2R1 to a 2300 series airfoil?
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Stoney,
I't doesn't have much relevance to the discussion I know... but just for clarification,
I don't know for all airplanes, but I know that the 109 Fs and later did not have the same 2300 series airfoil that the 109Es had.
109E arifoil - NACA 2314(root)- NACA 2310(tip)
109F+ airfoil - NACA 2R1 14.2(root) - NACA 2R1 11.35(tip)
how close is a NACA 2R1 to a 2300 series airfoil?
Dave Lednicer's list: http://www.ae.illinois.edu/m-selig/ads/aircraft.html shows all models of the 109 using the NACA 2R1 series airfoils. An older version of this list had the 2300 series listed for the early 109's, but it appears to have been changed. The 2R1, from what I've been able to find online, was a modified Clark Y airfoil that reflexed the lower surface right before the trailing edge. Apparently this was done to minimize the pitching moment. I still cannot find a profile for it, but one message board I found stated that it is in the Profili database, if anyone has access to that program.
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Here's a chart. 8 degrees AOA corresponds to a lift co of about .75 with flaps up.
(http://beta.hitechcreations.com/pyro/p51ddrag.jpg)
Many thanks, Pyro.
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I also do not believe the B-24 had any laminar flow (beyond 2-3mm). Please site anything other than wikipedia if you have it.
http://www.dreesecode.com/primer/airfoil5.htm
It shows the shape of the Davis airfoil vs. the NACA 6-series laminar-flow airfoil:
(http://www.dreesecode.com/primer/p5_f001.jpg)
Also, this, showing the Davis airfoil. Not as good as the P-51 airfoil, but quite good for its day:
(http://www.dreesecode.com/primer/p5_f002.jpg)
Of course, with dirt, dents, scratches, propwash, design variation, wing vibration, etc., the laminar flow wouldn't be at 20% of chord anymore. For all I know, maybe it is, in real-world application on real, in-service B-24's, down to something quite low. But -- it looks like Davis did make a very nice airfoil shape back in the days before widespread laminar-flow airfoils.
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But -- it looks like Davis did make a very nice airfoil shape back in the days before widespread laminar-flow airfoils.
No doubt, especially considering the knowledge base and technology of the day. I didn't mean for any of my explanations to imply otherwise--just wanted to tune up the terminology. :)
It actually looks very close to a NACA 4418...
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OK, since the bet didn't go through, I guess now I will go through my reasoning. (Altough I'm still willing to take the bet! :) )
Here is a graph of a NACA 6-series airfoil, actual test data. The P-51 wing is based on a 6-series airfoil. So while I didn't have the data for the actual P-51 itself (although Pyro provided it -- thank you, Pyro), I figured this would be quite close.
From Aerodynamics, Aeronautics, and Flight Mechanics, by McCormick, p. 69:
(http://www.electraforge.com/brooke/misc/aces_high/6seriesGraphs-4.png)
An 8 degree angle of attack of the airfoil (which is a 7 degree angle of attack of the P-51, since wings are mounted with an angle of incedence at root of about 1 degree, so I could really be looking at a 9 degree AoA for the airfoil, but let's use 8 for sake of argument) equates to a Ct of about 1.0.
At a Ct of about 1.0, clearly the drag is higher than at an AoA of zero degrees. This is true both for the "ideal condition" airfoils (very clean, no scratches, no dirt etc.) and certainly for the airfoil with "standard roughness" (just painted -- still no bugs, dents, scratches, etc.). Note two things. First is that "ideal condition" airfoil has the so called "drag bucket," which is the low-drag, laminar flow characteristic Chalenge was correctly indicating in principle, but it is not close to encompassing 8 degrees AoA -- it goes out to only about 2-4 degrees AoA. Second is that "standard roughness" totally eliminates all of that, and you have the more-typical parabolic shape to Cd vs. Ct, where you are quite guaranteed (as Stoney was pointing out from the usual modelling equations) to have higher Cd at a higher Ct.
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Having the 51 radiator scoop were it is should help with reducing air turbulence from entering the intake. The scoop design looks similar to the F-16 Falcon. I've read/heard the air turbulence is pretty bad along an airplanes skin and is especially bad for air intakes.
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...
An 8 degree angle of attack of the airfoil (which is a 7 degree angle of attack of the P-51, since wings are mounted with an angle of incedence at root of about 1 degree, so I could really be looking at a 9 degree AoA for the airfoil, but let's use 8 for sake of argument) equates to a Ct of about 1.0.
...
This is confusing me. Isn't AOA always determined by the wing regardless of incidence?
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This is confusing me. Isn't AOA always determined by the wing regardless of incidence?
The angle between the chord line of the wing and the relative wind = angle of attack. Therefore, it doesn't matter what the angle of incidence is.
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First is that "ideal condition" airfoil has the so called "drag bucket," which is the low-drag, laminar flow characteristic Chalenge was correctly indicating in principle, but it is not close to encompassing 8 degrees AoA -- it goes out to only about 2-4 degrees AoA. Second is that "standard roughness" totally eliminates all of that, and you have the more-typical parabolic shape to Cd vs. Ct, where you are quite guaranteed (as Stoney was pointing out from the usual modelling equations) to have higher Cd at a higher Ct.
Okay first things first...
Looking at your wind tunnel report (which is obvious thats what it is) I have to start off with the critique by saying (asking) where is the wind tunnel turbulence indicated? I dont see it and without that its not worth the graph.
Second I notice the airfoil is not a P-51 airfoil. It is similar to the P-51 tip section but it really isnt even close to that!
Here we are with the P-51D root and then tip followed by the P-51H root (which as I said was modified to assist with "Slipstream Contraction."
(http://i447.photobucket.com/albums/qq197/Chalenge08/p51droot.gif)
(http://i447.photobucket.com/albums/qq197/Chalenge08/p51dtip.gif)
(http://i447.photobucket.com/albums/qq197/Chalenge08/p51hroot.gif)
These same images and more data can be found at the UIUC Airfoil Coordinates Database.
I believe you can see the airfoils are quite different.
Now... back in 1959 a man by the name of August Raspet (PhD) at MU was conducting experiments with methods of flow analysis. In the glider world (soaring actually) Dr. Raspet is well thought of even now more than 50 years following his death (air accident during an experiment). He invented a technique that allowed experimenters to 'listen' to air flow over a give wing in flight. His method allowed one to clearly distinguish between laminar and turbulent flow in actual flight (not in a wind tunnel). He was particularly interested in airfoils that were of laminar design because it was his objective to eliminate (as much as possible) any form of drag. It is because of him that soaring pilots today have laminar foils on their gliders that actually experience laminar flow (like it or not).
Primarily his research focused upon foils that were laminar but not symmetrical. If you look at the P-51D you will see that as the wing transitions from root to tip the symmetry changes quite a bit. This change in symmetry has the effect of broadening the bucket. You will also discover that since the symmetry is not consistant that the wing must be analysed by local coefficients which was not easy then and isnt easy today.
However... take a close look at your standard roughness again and then consider what I just revealed to you. Not only is the graph going to sink on the scale of cd but it is also going to flatten out.
All the way up until 1979 (and perhaps beyond) the various agencies conducting airfoil analysis had a problem in that they could never correlate their data from one wind tunnel to the next. You have to be very careful what you accept and what you dont but certainly without published details on things like 'specific tunnel turbulence' you must be very careful.
Anyway... you can read more about this and other details about Raspet and the P-51D laminar research in the August 1960 issue of Soaring Magazine although it isnt really worth researching for the few details he gives.
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Well, then it is your lucky day. For while you are convinced that the data I posted is irrelevant (and presumably the same for the graph that Pyro posted), I think that it is spot on and am still willing to take the bet with you.
Are you willing to bet, and if so how much shall we bet?
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The angle between the chord line of the wing and the relative wind = angle of attack. Therefore, it doesn't matter what the angle of incidence is.
Angle of incendence does indeed matter when you say "This is why a P-51 can raise its nose as much as eight degrees (or a little more) and the drag remains the same". A P-51 with its nose raised eight degrees has its wing root at 9 degrees angle of attack. That was the point I was making.
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I would be willing to bet a large sum of money that a P-51 with wing at 8 degrees angle of attack has substantially more drag than a P-51 with wing at zero degrees angle of attack.
According to Pyros graph at a cl of 0.75 (8 degrees AOA) to cd is .008. At zero degrees AOA the cl is 0 and the cd is .00025 (approximately). This chart was produced in 1946 by calculation (apparently). There does not appear to be any relative published drag coefficient in any of the evidence you have that can be accurate.
However... you can still back out of your claim or more clearly define "substantially more drag."
I will tell you that the definition of "parasite drag" in the U.S. is commonly accepted as the total drag excluding vortex drag (your induced drag). In looking at Pyros chart I would have to say the chart defines parasitic drag as only the part of drag associated with surface friction/resistance with the air flow and does not mention profile drag (which as I understand things should double the drag coefficient since profile drag is usually the equal of or slightly greater than parasite drag). In other words the total drag would be 2 x .008 + vortex drag but in reality it would be more. The published zero-lift drag coefficient for the P-51 is .0163 and so please tell me what you think the total drag coefficient for 8 degrees angle of attack would be and explain how it is "substantially more drag."
Then you can explain why you think an airfoil that is nothing like the P-51s has any bearing on the subject.
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Chalenge, I am ready to wager with you on it.
We can put our money with a mutually trusted third party, draw the pay for three experts from the pot, each expert will render a decision on who is more correct, and the remainder of the pot goes to whoever gets the majority nod of the experts.
The bet will probably need to be about $2k each to cover cost of the experts and still have a profit left for the winner.
Will you take this bet?
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Angle of incendence does indeed matter when you say "This is why a P-51 can raise its nose as much as eight degrees (or a little more) and the drag remains the same". A P-51 with its nose raised eight degrees has its wing root at 9 degrees angle of attack. That was the point I was making.
I thought this originally started with a statement about "8 degrees angle of attack"?
I guess this isn't so much a discussion of the issue as it is an argument between you two. With that, I'll bid you both adieu.
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In looking at Pyros chart I would have to say the chart defines parasitic drag as only the part of drag associated with surface friction/resistance with the air flow and does not mention profile drag (which as I understand things should double the drag coefficient since profile drag is usually the equal of or slightly greater than parasite drag). In other words the total drag would be 2 x .008 + vortex drag but in reality it would be more. The published zero-lift drag coefficient for the P-51 is .0163 and so please tell me what you think the total drag coefficient for 8 degrees angle of attack would be and explain how it is "substantially more drag."
Then you can explain why you think an airfoil that is nothing like the P-51s has any bearing on the subject.
Actually, the 65215 is pretty close to the D model airfoil. Probably a 64215 would be closer, but the behavior of all the laminar airfoils plots out the same, so I think its a decent relative comparison.
The profile drag is shown on Pyro's chart--that's the type of drag being represented by the drag polar. There is no other drag represented except for the "standard roughness" graph, which introduces the skin friction component.
[EDIT]...after thinking further, even the standard roughness is still profile drag...
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Chalenge, I am ready to wager with you on it.
We can put our money with a mutually trusted third party, draw the pay for three experts from the pot, each expert will render a decision on who is more correct, and the remainder of the pot goes to whoever gets the majority nod of the experts.
The bet will probably need to be about $2k each to cover cost of the experts and still have a profit left for the winner.
Will you take this bet?
You'd make a bet with Paul Hinds? Good luck collecting.
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You'd make a bet with Paul Hinds? Good luck collecting.
You can be almost certain that one of Chalenge's experts will be an entomologist.
ack-ack
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I thought this originally started with a statement about "8 degrees angle of attack"?
I guess this isn't so much a discussion of the issue as it is an argument between you two. With that, I'll bid you both adieu.
It started with this statement: "This is why a P-51 can raise its nose as much as eight degrees (or a little more) and the drag remains the same," and whether or not that is correct is the topic of the discussion.
It's not that big a deal if one wants to substitute "raise the AoA of its wing as much as eight degrees" for "raise its nose as much as eight degrees," but the two are technically different by about one degree. That's all I was pointing out as one small part of my overall discussion. My feeling that the statement is not correct does not hinge on whether one is taking that to mean 8 deg. AoA or 9 deg. AoA.
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It started with this statement: "This is why a P-51 can raise its nose as much as eight degrees (or a little more) and the drag remains the same," and whether or not that is correct is the topic of the discussion.
It's not that big a deal if one wants to substitute "raise the AoA of its wing as much as eight degrees" for "raise its nose as much as eight degrees," but the two are technically different by about one degree. That's all I was pointing out as one small part of my overall discussion. My feeling that the statement is not correct does not hinge on whether one is taking that to mean 8 deg. AoA or 9 deg. AoA.
Sorry, I misunderstood then...
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Sorry, I misunderstood then...
No problem at all, and you are correct in pointing out that, for angle of attack of the wing, angle of incidence doesn't enter into it.
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You'd make a bet with Paul Hinds? Good luck collecting.
Karaya what is it with you and Mr Hinds? Did he send you to places you havent seen since? Grow up grow a pair and move on your gay bf is long gone.
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Actually, the 65215 is pretty close to the D model airfoil. Probably a 64215 would be closer, but the behavior of all the laminar airfoils plots out the same, so I think its a decent relative comparison.
The profile drag is shown on Pyro's chart--that's the type of drag being represented by the drag polar. There is no other drag represented except for the "standard roughness" graph, which introduces the skin friction component.
[EDIT]...after thinking further, even the standard roughness is still profile drag...
Okay Stoney if its 'pretty close' then I suppose 'substantially more drag' cant be verified.
The point is Brooke is trying to substantiate his argument with data that is not relevant to the P-51D. That leads me to believe that he either doesnt have verifiable evidence (which admittedly would be difficult to find) and/or he is incapable of working it out logically. Thats why he has turned to the betting approach.
My point is that even the evidence that I have seen from history (wind tunnel tests of the later 40s) are not quantifiable because of the tunnels having inherent turbulence. Some of the tunnels of the day had a turbulence factor of as much as 2.64 which requires not only a quantitative but statistical adjustment rendering any result useless.
Now if you want to look at data that is 'pretty close' then I suggest you do some research on the work of Eppler and Wortmann (pre-Raspet) where Eppler and Wortmann independently tested the NACA 65 415 with the Eppler 266 profile. The double bucket result is very interesting but again... 'pretty close' doesnt cut it.
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Thats why he has turned to the betting approach.
No, it's because of two reasons. First, is that I said I'd bet your statement was incorrect, initially along the lines of a figure of speech. But when you wrote "but your large sum probably equates to about $15," I thought "OK, let's go for it then." Second, no amount of discussion by me, no matter what data I find or what explanations I apply, is going to convince you. The only way to further progress is to get the opinion of a panel of undisputed experts and have them opine on it. But that likely will cost some money. Might as well have a way for whoever is wrong to foot the bill for that. My suggested wager covers both issues nicely.
You haven't accepted the wager, so I assume that your answer is "no," but that your way of delivering that answer is just not to say anything.
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Its not likely you have the ability to pay for the only kind of test that would settle it. As I already said... you posted one airfoil that is not the P-51. The P-51 airfoil is different at every station (rib) of the wing and so a single section isnt going to work either way. One of the reasons the 1946 data is 'calculated' is because the individual cl(s) need to be individually tested (probably with a very well designed 'drag rake') and then properly summated in a proper series. Modern tunnels have a turbulence factor of less than 1.06 but its very expensive to do this type of research and for the P-51 alone you will be using the tunnel for a week or more.
Even though I have the kind of money it takes Im not even going to pretend to be bold enough to put that on the line. If you are willing to do that your crazier than you look. :D
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Karaya what is it with you and Mr Hinds? Did he send you to places you havent seen since? Grow up grow a pair and move on your gay bf is long gone.
Even your former squaddies have come forward. I'm happily married. I realize you're under some stress so if you need to "tap" the left foot, leave me out of it.
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Okay Stoney if its 'pretty close' then I suppose 'substantially more drag' cant be verified.
The point is Brooke is trying to substantiate his argument with data that is not relevant to the P-51D. That leads me to believe that he either doesnt have verifiable evidence (which admittedly would be difficult to find) and/or he is incapable of working it out logically. Thats why he has turned to the betting approach.
My point is that even the evidence that I have seen from history (wind tunnel tests of the later 40s) are not quantifiable because of the tunnels having inherent turbulence. Some of the tunnels of the day had a turbulence factor of as much as 2.64 which requires not only a quantitative but statistical adjustment rendering any result useless.
Now if you want to look at data that is 'pretty close' then I suggest you do some research on the work of Eppler and Wortmann (pre-Raspet) where Eppler and Wortmann independently tested the NACA 65 415 with the Eppler 266 profile. The double bucket result is very interesting but again... 'pretty close' doesnt cut it.
You are exasperating... I've never read any of Eppler and Wortmann's work, admittedly. I know Eppler was part of the NLF series airfoil designs. But I have read Theory of Wing Sections from stem to stern. You look at enough drag polars and you can see trends. I've plotted and compared the 45-100 root airfoil plot with the 64218 plot and they're nearly identical. At the tip, the 64212 is almost identical. At the P-51D MAC, the 64215 plot is almost identical. This is assuming that the plot that you got from the UIUC airfoils database for the 45-100 is dead-on accurate. They are close enough for any reasonable person to use for comparisons. The differences? The 45-100 point of maximum thickness is at about 37%, same as the 64000 series. There is a bit of a variation in the mean lines used for the two, as NACA used the 1.0 mean line versus a shallower mean line used by North American. I don't know which because its impossible to tell visually. What does a shallower mean line mean? That the P-51D pitching moment would be a bit more benign, and add just a hair more profile drag. I'm talking variations of fractions of percent here. If I had time (probably take 6-8 hours of work), I could plot a 64418 a=.5, 64415 a=.5, 64412 a=.5 and see how much closer they would be, but I'm not.
I love these discussions when they're educational and enjoy the back and forth. This one could be, but isn't anymore.
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Its not likely you have the ability to pay
Even though I have the kind of money it takes
why do you constantly try to make everyone think you are better than them? get over yourself, cupcake. did u not see Pyro's post above?! my moneys on Brooke. but of course I don't have nearly the $$$ you do, as you so eloquently stated
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Even your former squaddies have come forward. I'm happily married. I realize you're under some stress so if you need to "tap" the left foot, leave me out of it.
If your talking about Uptown your talking about nothing. Otherwise your confused which is obvious.
Stoney: Your right end of thread. Brooke obviously had nothing to begin with except a desire to argue.
These other moonbats never have anything to add to any thread.
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If your talking about Uptown your talking about nothing. Otherwise your confused which is obvious.
Stoney: Your right end of thread. Brooke obviously had nothing to begin with except a desire to argue.
These other moonbats never have anything to add to any thread.
Sorry, you're incorrect on your "assumption". Keep guessing though.
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I considered posting a unified aero oriented view of the topic on the importance of paying attention to the boundary layer with the P-51 as the case study but given the direction the thread is going I decided against it. That being said here are a few points I thought I would point out.
1) Eppler and Wortmann are red herrings for all sorts of reasons, one being that the reynolds numbers we deal with in WW2 fighter planes don't apply.
2) All wings have both laminar and turbulent flows in the boundary layer. The question is where the transition occurs on the chord of the wing between laminar and turbulent flows. Conventional wings have the transition point between 5%-20% of the chord from the leading edge. Low drag laminar flow wings were designed to try and move the transition point further back along the wing compared to conventional airfoils increasing the length of the laminar runs which reduces skin friction drag.
3) Though true that the P-51 wing in "service condition" didn't produce the longer laminar flows, this could be rectified by polishing the wing and smoothing out surface waviness which was the main culprit. Here's the data from the NACA tests on the XP-51 that show's this. There are also other data points from other NACA reports that demonstrate this with other low-drag wings.
XP-51 Surface Waviness – service condition vs. smoothing:
(http://brauncomustangs.org/upload/files/p-51surface.jpg)
XP-51 Drag Polar – various surface finishes of the wing:
(http://brauncomustangs.org/upload/files/p51-drag.jpg)
4) Whether smoothed out or not the "laminar bucket" does not exist for the drag polar of the complete P-51 airplane. Pyro's chart show's this. If this isn't enough the following charts from P-51 flights tests, wind tunnel tests, and modern CFD analysis show this as well.
P-51B Flight Test Drag Polars:
(http://brauncomustangs.org/upload/files/p51dragpolar-flight.jpg)
P-51B Wind Tunnel Drag Polar:
(http://brauncomustangs.org/upload/files/p51dragpolar-wt.jpg)
P-51B Modern CFD Drag Polar:
(http://brauncomustangs.org/upload/files/p51dragpolar-cfd.jpg)
5) Parasite drag by definition includes both the skin friction drag due to boundary layer fluid shear stress and pressure / profile drag due to boundary layer separation from viscous effects. Pyro's chart is the increment in parasite drag (skin friction & pressure/profile drag) with lift coefficient for the P-51.
6) The data that Brooke pointed out from McCormick is data on NACA 6A series airfoils conducted in the low-turbulence wind tunnel at Langley (NACA Report No. 903).
Hopefully that clears a few things up.
BTW - Brooke would win the wager ;).
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Thanks for the post, dtango.
Well, there you have it.
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i don't know much about physics, but i do know that a wing at 8 degrees AoA is going to produce more drag than a wing at 0 degrees AoA. it's preschool simple. to say otherwise is the same as saying the square peg goes in the round hole. geez, i figured that one out by 7th grade :rolleyes:
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why do you constantly try to make everyone think you are better than them? get over yourself, cupcake. did u not see Pyro's post above?! my moneys on Brooke. but of course I don't have nearly the $$$ you do, as you so eloquently stated
He also has the aircraft it takes, a vintage P51D, that is currently sitting in a hangar in Nantucket.
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BTW - Brooke would win the wager ;).
No... I tried to get him to clarify on what he would say 'significant' was which he didnt.
My point on Eppler and Wortmann was that it was the driving factor on Raspet looking into assymetrical laminars which is what the pony has I believe. Raspet himself is the one that suggested the range of eight degrees AOA. I know I know the closer examination would show minute changes in cd but thats the point of the double bucket after all. Raspet defined the Mustang as partially turbulent at all times but also laminar (upper or lower surface) in most situations ruling out the case of stalls.
I am well aware of the surface finish research. Thankfully we have quite different materials today.
As to your definition of 'parasite drag' I know it is common for Americans to believe they (we) have established the universal rule but in this case it is not true. The European (european union) and I believe Australian/New Zealand definition differs.
as now generally accepted, any or all of the drag forces acting on an aircraft that are not formed in the production of lift. In subsonic flow, any or all of the drag forces acting on an aircraft exclusive of the induced drag b) in some contexts, especially in the older literature, any or all of the drag forces from parts of the aircraft that do not contribute to the lift. In sense (b), the parasite drag does not include 'profile drag'.
And Brookes data is still not a Pony airfoil. 'Close enough' or not.
One might ask why it is IrishOne and Grizz decided to show their ugly butts and attempt to nail the lid shut on this one. I know neither one of them could possibly keep up with the conversation. :rolleyes:
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it's preschool simple. to say otherwise is the same as saying the square peg goes in the round hole.
Life wisdom: The square peg goes in the round hole if the hole radius is sqrt(2) times larger than the square's side length. :old:
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See Rules #4, #6, #2
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Raspet is famous for his contributions to laminar flow wings for sailplanes. The reason that longer laminar flows are achievable for gliders is because they operate at lower reynolds numbers compared to powered aircraft, most notably our WW2 fighters. However at higher reynolds numbers the boundary layer becomes even more unstable and it is much more difficult to obtain longer laminar runs.
Increasing reynolds number decreases boundary layer thickness which makes the boundary layer much more susceptible to all sorts of things that trip laminar flows to turbulent like local disturbances caused by surface waviness from the manufacturing process. For pre-1950's construction it was downright impossible to obtain long laminar flows without the aforementioned refinishing and smoothing. The XP-51 with standard factory finishing only acheived laminar flow for 15% of the chord. Every credible aero report and aerodynamic text I've seen unequivocally acknowledge the dismal record of obtaining laminar flows on low-drag wings for higher reynolds flight and it's only been in the last 30 years or so that this has changed.
As to the whole 8 degrees aoa issue, even if you could achieve the laminar drag bucket for the wing as evidenced in all the drag polars for the P-51B I posted you wouldn't even see it because as shown other pressure drag from the entire P-51 dominates with increasing aoa which results in the traditional parabolic drag polar.
As to the definition of parasite drag, well we could quibble what profile drag is. However drag not associated with lift is definitely a result of viscous forces- fundamentally skin friction due to shear stress and pressure drag due to separation. As to your definintion of parasite drag, hmmm I think I'll settle with what's taught in aeronautics classes around the world and in the aero industry vs. from webster's online dictionary ;).
Tango
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I tried to get him to clarify on what he would say 'significant' was which he didnt.
OK, let's say "substantially more drag" is at least 50% more drag.
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I know neither one of them could possibly keep up with the conversation. :rolleyes:
:lol
You're right, I don't know much about laminar flows, turbulent flows, reynolds numbers, Cd values,etc. pertaining to aircraft. Just the basic theory that I learned in my fluid mechanics civil engineering course years ago.
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Regarding this (the original reason for the discussion):
Also... the laminar airfoil of the P-51 eliminates the increase of the coefficient of drag for the majority of the effective AOA. What that means is that as a pilot increases the AOA the coefficient of lift increases but the coefficient of drag remains constant for 65% (approximately) of the effective AOA. This is why a P-51 can raise its nose as much as eight degrees (or a little more) and the drag remains the same as if the wing were level and the P-51 can use its E to zoom to great advantage.
It dawned on me today why you might be thinking that. I wondered if you might be considering the section drag coefficient as being the same as the drag coefficient of the wing. Section drag coefficient is drag of the airfoil -- i.e., a wing of infinite aspect ratio. Once you have a wing of finite aspect ratio, it necessarily (even if the wing is in all other respects idealized) introduces induced drag that increases with lift (i.e, with angle of attack), and it is appreciable. There is no way around it -- it is a physical result of a wing of finite wingspan, whether the airfoil is laminar flow or not.
Let C_D = coefficient of drag.
For an idealized airfoil (infinite aspect ratio), C_D = c_d = constant, where c_d is the section drag coefficient (also "parasitic drag coefficient", "profile drag coefficient", etc. for the airfoil, and using various symbols, such as "C_D_0", "C_D_min", "c_d", etc. depending on reference). In the graphs I posted, it's called "section lift coefficient". In Pyro's graph, "parasite drag coefficient". In dtango's graph, "profile drag coefficient".
For an idealized wing producing constant value of downwash across the span (elliptical lift distribution is the optimal form here), C_D = c_d + C_D_i, where C_D_i = coefficient of induced drag = C_L^2 / (pi * AR), C_L is the coefficient of lift, pi is pi, and AR is the wing's apsect ratio. This is the way it is once one has a wing that is no longer infinite wingspan -- its a physical derivation that lift must induce drag of this form.
For wings that diverge from the optimal elliptical lift distribution, C_D_i = C_L^2 / (pi * AR) * (1 + delta), where delta is typcially small.
For real (non-idealized) airfoils, c_d does vary with C_L. For non-laminar-flow airfoils, c_d or section drag coefficient is roughly parabolic with C_L. For laminar-flow airfoils (ones where laminar flow happens for an appreciable amount of the chord), c_d is typcially better than for non-laminar-flow airfoils, and it can have a "drag bucket", where c_d is low and nearly constant for some amount of C_L before laminar flow is eventually lost and c_d goes back to being parabolic vs. C_L.
However, c_d is much less than induced drag beyond small angles of attack, so that isn't even needed to conclude major things about drag at 8 deg AoA. Thus, while I believe that for the P-51's airfoil c_d (or section drag coefficient) at 8 degress AoA is much greater than at zero deg. AoA (and you disagree), that doesn't even matter to the discussion of whether the P-51's wing has more drag at 8 deg. AoA than at zero AoA. c_d (i.e., C_D at zero lift) is on order 0.001 to 0.01 or so. C_D_i is much larger.
Let's look at some numbers. Let's (contrary to reality but for the sake of argument) take the P-51's airfoil to be 100% laminar flow and 100% the same as a perfect, idealized airfoil so that it's section drag coefficient is totally independent of C_L (or thus AoA) and equal to c_d_min. From the graphs I posted (and numerous other graphs of airfoil data), I'd guess that c_d_min for the P-51's airfoil is somewhere around 0.004, but also for the sake of argument, let's take it to be anything you want in the range of 0.001 to 0.01. Now let's look at C_D_i = C_L^2 / (pi * AR) (not even putting in any value for delta -- for the sake of argument letting the P-51 have an optimal elliptical lift distribution). The P-51's AR is about 5.8, so C_D_i = 0.0549 * C_L^2. At zero lift, C_D = c_d_min, which equals, say, 0.01. At 8 deg. AoA, where C_L is about 1.0, then C_D = 0.01 + 0.0549 = 0.0649 -- more than six times higher than C_D at zero lift. Or, if anyone wants to dispute that 8 deg AoA isn't a C_L of 1.0, fine -- put in something unrealistically low for the sake of argument -- say, 0.5. At C_L = 0.5, C_D = 0.01 + 0.0137, and C_D at 8 deg. AoA is still more than two times what it is at zero lift. In all cases, I would call that "substantially more".
References for the above, so that (although it is standard material in aerodynamics books), I can't be accused of making it up or misderiving it.
Airplane Performance, Stability, and Control, by Perkins and Hage
Aerodynamics, Aeronautics, and Flight Mechanics, by McCormick
Fundamentals of Flight, by Shevell
Theory of Flight, by Von Mises
Introduction to Flight, by Anderson
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Chalenge, any thoughts on the above?
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See Rules #2, #4
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See Rules #2. #4
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See Rules #2, #4
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On the discussion of laminar flow, how did the Mosquito stack up in comparison to the P51s "laminar flow design"?
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On the discussion of laminar flow, how did the Mosquito stack up in comparison to the P51s "laminar flow design"?
It used a "turbulent" RAF-34, according to Dave Lednicer. Higher drag, possibly higher Clmax, probably more docile stall characteristics.
EDIT: This is just a comparison of the airfoils, not the aircraft. The wing and fuselage design of the Mossie, and the resulting aerodynamic efficiency, are another matter all together...
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Well, Chalenge as been on the board lately, so I guess his lack of posting is just that he's got nothing to post that is in disagreement with the latest.
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Well, Chalenge as been on the board lately, so I guess his lack of posting is just that he's got nothing to post that is in disagreement with the latest.
This is nothing new. He will never admit being wrong. The best you will get out of him is a disappearing act.
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This is nothing new. He will never admit being wrong. The best you will get out of him is a disappearing act.
Correct!