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Author Topic: P47 vs 190  (Read 7063 times)
SectorNine50
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« Reply #105 on: March 04, 2011, 04:15:09 PM »

What is the "camber on the pitching moment of airfoils?"  I'm trying to picture what a wing with more camber looks like vs. one with less.
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« Reply #106 on: March 04, 2011, 04:19:29 PM »

What is the "camber on the pitching moment of airfoils?"  I'm trying to picture what a wing with more camber looks like vs. one with less.

Try this:  http://www.centennialofflight.gov/essay/Theories_of_Flight/Airfoils/TH13G2.htm
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« Reply #107 on: March 04, 2011, 04:19:38 PM »

Stoney,
That was very informative, thank you.


So if I understand the diagrams correctly, the 2300 should allow for a higher angle of attack (~17 degrees) over the ~15 on the 2200 series, correct?

This would suggest that then that a FW 190 should be able to pull a bit steeper than clark-y which is roughly a 2200 ( NACA 2R1 14.2). Stoney, Also, the 109s clark Y airfoil was added in ~1941, previous 109s had  NACA 2314 air foils, aka f-k models had clark-y.
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« Reply #108 on: March 04, 2011, 04:31:09 PM »

Stoney  Maybe a stupid question, but hopefully simple enough to answer - is both increased camber (all else being equal) and low wing flaps causing nose pitch down a coincidence?
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Stoney
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« Reply #109 on: March 04, 2011, 04:37:39 PM »

Stoney,
That was very informative, thank you.


So if I understand the diagrams correctly, the 2300 should allow for a higher angle of attack (~17 degrees) over the ~15 on the 2200 series, correct?

This would suggest that then that a FW 190 should be able to pull a bit steeper than clark-y which is roughly a 2200 ( NACA 2R1 14.2). Stoney, Also, the 109s clark Y airfoil was added in ~1941, previous 109s had  NACA 2314 air foils, aka f-k models had clark-y.


No, don't make that assumption based purely on those graphs.  Those graphs are snapshots of certain flight regimes built to demonstrate my argument.  First, the chart with the 23000 series on it is showing the same airfoil at two different thicknesses (9% and 15% which were common tip/root thicknesses used on most designs that used the 23000 series) using the same Reynold's number.  The 2200 series chart shows the same 9% thickness airfoil at two different Reynold's Numbers.  To do an accurate comparison between the two different airfoils, you would need to use the same thicknesses and same Reynold's number.

Thanks for the correction on the Clark Y/109...


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« Reply #110 on: March 04, 2011, 04:46:30 PM »

Stoney  Maybe a stupid question, but hopefully simple enough to answer - is both increased camber (all else being equal) and low wing flaps causing nose pitch down a coincidence?

No it is not.  Whenever flaps are dropped, you increase the camber of flapped portion of the wing considerably.  It increases the Clmax of that portion of the wing as a result, which allows for slower landing speeds.  Now, some aircraft experience differences in the amount of nose-down pitching moment with flap use, based on all the other moments on the aircraft, but to make a general, simplistic statement, dropping flaps will make the nose pitch down.

Just be careful not to expect this to apply to all aircraft.  I don't have a lot of time in Cessna 172's, so I can't remember, but the last time we debated this, someone that does stated that the 172 exhibits a nose-up pitching moment with flaps.  I know in my Grumman, your hand goes directly from the flap switch to the trim wheel to introduce a ton of nose-up trim to counteract the increase in pitching moment.  The Grumman is low-winged, and has a 64415 airfoil that has a very high pitching moment anyway.  So, pitching moment of the aircraft with flap use is up for debate.  Increasing the nose-down pitching moment of the airfoil with increased camber is not.
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« Reply #111 on: March 04, 2011, 04:49:43 PM »

Thanks again Stoney.
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SectorNine50
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« Reply #112 on: March 04, 2011, 05:36:04 PM »


Thanks much, all makes sense.
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« Reply #113 on: March 04, 2011, 07:30:09 PM »

Well, I don't know so much about the Clark Y and I've never been able to find a lot of data about it.  The 23000 series NACA airfoils, on the other hand, were the most prevalent and widely used airfoil in the history of aviation probably.  Many of the WWII fighter aircraft in-game shared the 23000 series airfoils, or close variations.

Stoney the Clark Y and the Gottingen 398 were early efficient airfoils.  When the two were compared it was found that when camber was removed both airfoils exhibited the same characteristics were observed for similar maximum thicckness.  IIRCthe four digit NACA series sprang from these two airfoils

The development of the 23000 series airfoil came out of the initial research of NACA (later to be known as NASA) and a lot of their wind tunnel testing in the 20's and 30's.  As more and more planes switched from fabric (which didn't hold a true airfoil form in flight necessarily) to metal (which did), designers began to realize certain performance aspects of the airfoils could be improved.  Now, remember that the Wright brothers had only flown 27 years prior to 1930, and a ton of the stuff we now know about aerodynamics was then still unknown.  One aspect of airfoils that designers did know a lot about was the effect of increased camber on the pitching moment of airfoils.  Camber is used to increase the Clmax of an airfoil, i.e. more camber = more lift given everything else the same.  When camber goes up, the airfoil has a nose down pitching moment.  When camber of some airfoils gets high enough to make the airfoil a very attractive choice from a lift perspective, it becomes a problem because a high enough nose-down pitching moment is introduced to create other problems.


This other problem is trim drag, or the drag created by the horizontal stabilizer counter-acting the nose down pitching moment of the wing.  When the pitching moment goes up, the trim drag goes up.

The Trim Drag increase occurs becaue the chord line angle between the leading edge of the horizontal stabilizer to the trailing edge of the elevator changes with the deflection of the elevator to assist in change of pitching moment of the aircraft - simialr ro deplpoying flaps. With 'increase in relative AoA comes an associated increase in drag - both Induced and parasite, of the horizontal stabilizer/elevator system'  

During this early period of airfoil design, minimizing pitching moment was seen as one of the most important characteristics of the airfoil chosen.  NACA decided that if the airfoil was designed properly, pitching moment could be minimized.  Their best result was the 23000 series of airfoils.  It had very low profile drag, and a very low pitching moment compared to some of the other airfoils that had been used previously.  As a result, it rapidly became the flavor of the month for airfoils, so-to-speak, and was widely accepted as a high-performance airfoil.  As a result, it was used in almost all of the later model WWII fighter aircraft like the F4U, F6F, P-47, FW-190, (Lednicer also lists the Ki-48 and N1K2 and some of the Russian fighters like the La-5/7).  The 23000 series can still be seen today in the Beechcraft Bonanza, and even in HiTech's RV-8, among others.

The problem is that (according to Harry Riblett) in changing the airfoil shape to minimize the pitching moment, NACA gave the 23000 series a very sharp leading edge profile on the top of the airfoil, and flattened the lower leading edge of the airfoil.  

Stoney I am confused on this one. by definition the 23000 series has a .02 radius (of chord), perfectly round at LE - why 'flat on bottom, particularly on a symmetrical aircfoils

This characteristic created a condition where, at the stall, airflow is disrupted further forward along the top edge of the airfoil than when compared to its contemporary airfoils.  This created very sharp stall characteristics, which were even more pronounced when airfoil thicknesses were thinner (see graph below).

Now, typically, the Clmax of an airfoil increases with thickness.  It also increases as the Reynolds number (a number that represents a relationship to the velocity X chord length of an airfoil) increases.  These two images are plots I made on XFoil a few years ago.  On the 23000 series comparison, you can see the impact of thickness on Clmax, and on the comparison of the 2200 series, you can see the impact of Reynolds number on Clmax.

<Quoted Image Removed>

<Quoted Image Removed>

Most WWII fighter aircraft used both planform taper (meaning that the wingtip had a shorter chord than the wing root) and also used thickness taper (meaning the wingtip had a thinner airfoil than the wing root).  Given the graphs I posted above, you can see that they basically designed wings that were going to have a tendency to tip stall at high Angle of Attack conditions, i.e. the wing tip stalls while the wing root is still producing lift.  At the same airspeed, the wing tip's thinner airfoil stalls sooner and more sharply, and the wing tip's lower reynold's number (since the wing tip chord is shorter than the wing root) is lower. The problem is that the ailerons, which control roll, are located on the wing tip.  If you get into a tip stall condition, there will be insufficient flow over the ailerons to continue to control the aircraft in the roll axis.  This is why when we stall fight aircraft in AH2, the planes want to roll over on their backs when we stall them, regardless of aileron input.  All of the other resultant roll moments (engine torque being a large one since we're almost always at full power in a stall fight) have more force than what the ailerons can counteract.

Now, why would they design the wings this way?  Lower drag, primarily, as profile drag decreases with airfoil thickness.  Ignorance would be another reason.  These guys were designing aircraft with nothing more than a plotting board and a slide rule.  Much of what we no know from CFD analysis was unknown then.  Also, they all thought, and many designers today still do think, that they could counteract all these tip stall tendencies with twist.  Unfortunately, most of the time, all the twist did was add more drag, while still not providing enough of a difference to counteract the tip stall tendencies they designed in with their combination of airfoil thickness and planform taper.  

Stoney - one reason for both leading edge twist and a planform taper was to drive the lift distribution inboard to more closely approximate an elliptical lift distribution. With twist however, you also increased drag due to lift as one of the penalties.  Both the 51 and the Spit had twist all the way out to the tip - the 190 only to about 80% span - why? who knows?

Ironically, the Me109, having been designed much earlier than some of its wartime counterparts, used leading edge slats to combat these tendencies, and were largely successful at controlling them.  While I've not done the analysis in XFoil, I would assume that the Clark Y airfoil also had better stall characteristics than the 23000 series.

So, I don't know if I answered your question entirely, but hopefully this gives you a start on understanding some of the issues.



/quote]
That was an excellent summary
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« Reply #114 on: March 04, 2011, 07:42:40 PM »

http://www.youtube.com/watch?v=kz8QsBxcOMw
The fight at 0:30 between a 190 A7 (basically an A8 with less boost) and a 47Dxx
Notice how the 190 follows the Tbolt thru several tight turns, pulling lead several times, even after several rotations, the Tbolt simply cant turn tight enough to shake the 190.
Thoughts?

edit  the 47 looks to be a razorback model, which is the best turning jug


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Debrody
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« Reply #115 on: March 05, 2011, 02:00:05 AM »

Thank you Stoney    salute
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Stoney
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« Reply #116 on: March 05, 2011, 03:19:12 AM »

Stoney the Clark Y and the Gottingen 398 were early efficient airfoils.  When the two were compared it was found that when camber was removed both airfoils exhibited the same characteristics were observed for similar maximum thicckness.  IIRCthe four digit NACA series sprang from these two airfoils

Perhaps.  I'd have to dig through my copy of Abbot and Doenhoff to check this, but good enough for me.  Like I said, I don't know much about the Clark Y...

Quote
The Trim Drag increase occurs becaue the chord line angle between the leading edge of the horizontal stabilizer to the trailing edge of the elevator changes with the deflection of the elevator to assist in change of pitching moment of the aircraft - simialr ro deplpoying flaps. With 'increase in relative AoA comes an associated increase in drag - both Induced and parasite, of the horizontal stabilizer/elevator system'

Yes.  I sort of skimmed the wave tops on the trim drag issue to keep from getting bogged down in the theory.

Quote
Stoney I am confused on this one. by definition the 23000 series has a .02 radius (of chord), perfectly round at LE - why 'flat on bottom, particularly on a symmetrical aircfoils

I was probably a little too vague with this.  I didn't have my copy of Ribblet's book at hand--its packed up--so I sort of swagged the description of this.  In the interest of full disclosure, Lednicer disagrees with some of Ribblet's conclusions in a response he wrote criticizing Ribblet's work, but given the well documented stall behavior of the 23000 series airfoil, they did something to it.  I don't understand it all myself, but what Ribblet postulated was the method with which they minimized the pitching moment of the airfoil modified the forward part of the bottom airfoil surface.  Not necessarily changing the leading edge radius, but the leading edge profile beyond the radius portion.  Something about how they flattened the area directly behind the radius on the bottom of the airfoil interferes with normal flow at very high AoA.  I'll see if I can't find the book and post an excerpt straight from the book.

In the meantime, looking at the top image on this page of a 23015, you can see the "flattened" portion of the airfoil lower surface, directly behind the leading edge out to about 25% chord. 

http://www.laboratoridenvol.com/info/tech/perfils.en.html

This is the portion of the airfoil that Ribblet argues "decambers" the forward portion of the airfoil, thus causing the very sharp stall characteristics that are exacerbated in the thinner 23000 series airfoils.  The 23015 doesn't have a very sharp stall, but if you look at the 23009 (which was the thickness used on the wing tips of most 23000 airfoil aircraft, it gets very severe.  I can't imagine a worse place to have a portion of the wing stall with very little warning.

Quote
Stoney - one reason for both leading edge twist and a planform taper was to drive the lift distribution inboard to more closely approximate an elliptical lift distribution. With twist however, you also increased drag due to lift as one of the penalties.  Both the 51 and the Spit had twist all the way out to the tip - the 190 only to about 80% span - why? who knows?

Again, I kind of skimmed the wave tops to avoid some of the complexities of planform taper, its advantages and disadvantages, etc.  For the most part now, designers use twist to combat tip stall tendencies, as they've determined that elliptical lift distribution can be most easily achieved through planform taper only.  Obviously some still disagree today.  On my Formula 1 design, I was using zero twist, and a 45% taper ratio to get close.  I planned to use a 15% thickness airfoil through the whole wing to avoid thickness taper.  I even considered using a 17% thickness at the wingtip before I decided to just keep the same thickness for construction simplicity.  Further, I was using a lower aspect ratio wing (for a lot of reasons), which made the chord at the wingtip long enough to have a decent Reynold's Number at landing speeds.  Another part of this I didn't touch on was that several WWII fighter planforms, especially the Russians', introduced sweep into the wing as well, which reduces Clmax, but again, I didn't want to get into that.
[/quote]
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« Reply #117 on: March 05, 2011, 07:05:07 AM »

Stoney - We are pretty much in violent agreement.  One thing I would pose is the original concept of CMac (IIRC from 45 years ago of 'theory of Aero).  I don't have a picture but if we were to plot out the pressure distribution 360 degrees around an airfoil in lift, then integrate the top surface chord-wise, from the LE stagnation point to the trailing edge we would get our force vector on the top surface. Perform the same integration on the lower surface to obtain the similar - but not same magnitude nor the focal point of action- for the bottom surface.

IIRC, the resolution of the two force vectors about the leading edge is CMle.  We then resolve the 'moment obtained thusly' and transpose both the resultant force vector and the Moment to the aerodynamic center of the wing (e.g. the 1/4 chord point) to obtain L and CMac.

So, futzing with the leading edge geometry of say a 23012 airfoil should result in two things. A dash number to explain the difference in the 'new 23012' airfoil geometry like a leading edge radius or sharpness distinction, and a new and different set of wind tunnel data to plot the differences in Cl, Cd and CMac as well as notes to alert the designer to nasty departure characteristics (hmm that would be three things?)

I dropped into another site on a discussion and noticed that the Do 335 had this airfoil with a three digit post series and I had not seen that before

For a two digit modification after the dash number - the first digit represents deviation from 'normal' leading edge radius (where 6=normal, 0=sharp) and the second indicates the position of the maximum thickness in percent chord. So the 23018-63 would be same as 23018 except the max t/c moved aft from 15 % to 30% of chord - more closely approaching 65 series laminar flow wings which were in the 40% range.

Having said this I have never seen any dash number after the Fw 190 airfoild series so I really don't know.
 
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« Reply #118 on: March 05, 2011, 07:44:04 AM »

A looooong looooong time ago the 190a5 used to be able to beat up on the P-47s in a knife fight. 

Then the code governing airflow and flaps were changed, and all the 190s went from being merely unmanuverable to being awful. The P-51s got hit with the nerf bat even harder... the P-51D went from being a roughly even match for a 109G-10 (todays K-4) to an easy kill for pretty much anything in a fight. I think the P-47 benefited from the change, but that may just be vis a vis the P-51 and 190 series. The F4Us and 109s were obvious winners - both series vaulted up to somewhere near the Spitfires in terms of 'turnyness'. Granted, the F4Us (with the exception of the -4) are still not hard to kill if you are in a Spit. They can turn with you, but you can get them turning and then go vertical and they won't have the power to keep up. 

The 109s on the other hand.... a 109F4 can hang with a pretty decent spit driver. The Spit 16 is still a better plane, but it isn't a blowout.

Now, things may have changed in the years since I stopped playing, but I'm pretty sure the flaps thing was the last major revision to the FMs.
i locve that song what should it be called?   "the day the 190s died"?
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« Reply #119 on: March 05, 2011, 09:48:36 AM »

Could just as easily be titled the day the runstangs died Smiley. If anything they took a bigger hit than the 190s. There weren't too many people who liked to knife fight in the 190s anyway (maybe 4-5). The P-51 going from a pretty good all-round fighter to being bait in any sort of turning fight made a much bigger difference in the grand scheme of things. You are talking about one of the most popular planes in the game here. I mean, the P-47s (any of them) absolutely eat the P-51s up in a turn fight.
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