Dora roll rate and turning rate, true to real-life data?

[KenobiOrder]

Yo-Yo charts and conclusions mirror what I wrote. It is up to the pilot, what profile he chooses and the skill in which he executes it are the deciding factor. The aircraft are equals in terms of sustained constant altitude turn performance.

 

As for the Clmax of the P-51...it is not far behind the FW190.

 

I get 1.51 for the P-51 vs 1.58 for the Dora.

 

Generally speaking A laminar flow airfoil does necessarily mean a low Clmax. In high angle of attack flight there is not an airfoil in existence that realizes any benefit from laminar flow. Laminar flow airfoils realize drag reduction benefits only at the bottom of their characteristic drag bucket at very low angles of attack in the vicinity of cruise flight and beyond.

 

Except they do not. Yo-Yo was clear in the turning thread that the P-51 out turns the FW190.

 

For starters, your assumption on ClMax is at odds with his:

 

"You do not take in account different CL especially at the speed the best turn ratio is. Mustang has better CL max at low M-numbers and much preferrable at medium."

[Hummingbird]

I have read it. And you are wrong.

 

Obviously you have not.

[Hummingbird]

How many times do I have to tell you the P51D didn't use NACA 66(1)-212 airfoil? The XP51F/G/J did.

The P51D used this airfoil:

http://airfoiltools.com/airfoil/details?airfoil=p51droot-il

Which had a 1.3-1.4 clmax.

 

You're the one who's incorrect here.

 

 

*sigh*

 

If you had cared to follow more closely you'd have noticed that I'm using two sets of data here actually, one is the page I provided earlier (which only shows results up to a Re of 1 mil sadly) and another is real test result figures from the 40's which provides figures at a Re of 6-9 mil, which is what we want really. Problem with 1940's tests is that they need to be side by side to be of accuracy, but the two I pulled up are infact from a side by side study so its good.

 

Furthermore the P-51 actually used an airfoil more similar to the NACA 66 designated the NAA/NACA 45-100, thus looking at the NACA 66's values is basically the same.

 

Also the real problem with the P-51s airfoil is not at the root, its as you travel outwards toward the tip where the Clmax drops drastically (You can even see that on the page I linked at a low Re number). By comparison the NACA 23XXX keeps a constant 1.60-1.55 clmax from root to tip.

[Hummingbird]

Except they do not. Yo-Yo was clear in the turning thread that the P-51 out turns the FW190.

 

For starters, your assumption on ClMax is at odds with his:

 

"You do not take in account different CL especially at the speed the best turn ratio is. Mustang has better CL max at low M-numbers and much preferrable at medium."

 

Which is just hopelessly false.

 

The NACA 23XXX airfoil provides a higher CLmax across the subsonic speedband, and I've got the documents to prove it too.

[KenobiOrder]

Which is just hopelessly false.

 

The NACA 23XXX airfoil provides a higher CLmax across the subsonic speedband, and I've got the documents to prove it too.

 

Perhaps you should re-read that thread, Yo-Yo explains quite clearly (about a billion times) why you are wrong regarding this. Bring all the documentation you want, we can all just refer back to that thread where you get soundly debunked. Honestly there is no point in debating this what-so-ever. The facts of this matter were explained several times ages ago. The Dora turns worse than the P-51, it has a worse CLmax than the P-51, particularly at the appropriate speeds for a sustained turn. Yo-Yo's FM in this regard is backed up by the fact that it is in perfect agreement with every sim going back a decade or so. No Dora out turns a P-51D at any comparable loading. It should be exceedingly clear at this point: especially since the pre-patch 2300BHP Dora was STILL incapable of turning with the P-51.

[Hummingbird]

NACA 23XXX type airfoil CLmax figures at an average Re of ~8.35 mill (1.6 to 1.68 from 9 to 15% thickness ratio):

[Hummingbird]

Perhaps you should re-read that thread, Yo-Yo explains quite clearly (about a billion times) why you are wrong regarding this. Bring all the documentation you want, we can all just refer back to that thread where you get soundly debunked. Honestly there is no point in debating this what-so-ever. The facts of this matter were explained several times ages ago. The Dora turns worse than the P-51, it has a worse CLmax than the P-51, particularly at the appropriate speeds for a sustained turn. Yo-Yo's FM in this regard is backed up by the fact that it is in perfect agreement with every sim going back a decade or so. No Dora out turns a P-51D at any comparable loading. It should be exceedingly clear at this point: especially since the pre-patch 2300BHP Dora was STILL incapable of turning with the P-51.

 

You are so wrong it hurts.

 

The P-51's airfoil features a lower CLmax, it's fact, deal with it.

 

I was never debunked and my points still stand firm.

 

Now let me repost Yo-Yo's chart for you:

http://forums.eagle.ru/attachment.php?attachmentid=116088&d=1428433410

 

Incase you're having trouble reading it the Dora on his chart actually beats out the mustang with a sustained 360 deg turn time of 21.6 sec vs the P-51's 21.8 sec, and this is despite using incorrect Clmax figures.

 

Sharpening the leading edge of an airofoil and moving back its point of max thickness will NOT yield good results for its low speed high AoA (and thus CLmax) capabilities, how Yo-Yo got this mixed up I have no clue

Edited April 8, 2015 by Hummingbird

[NineLine]

What is this thread about again? Whose argument can out turn the other person?

[KenobiOrder]

NACA 23XXX type airfoil CLmax figures at an average Re of ~8.35 mill (1.6 to 1.68 from 9 to 15% thickness ratio):

 

This is for only part of the foil.....not a wing. And therefore irrelevant. Go get a FW190 wing tunnel test of the entire wing etc and then you might actually have something.

[GrapeJam]

*sigh*

Furthermore the P-51 actually used an airfoil more similar to the NACA 66 designated the NAA/NACA 45-100, thus looking at the NACA 66's values is basically the same.

 

Oh really? I'd like to see the sources of that, meanwhile I've found this:

https://books.google.com.vn/books?id=fkat9yxMUGYC&pg=PA31&lpg=PA31&dq=naa/naca+45-100+airfoils&source=bl&ots=DoZR2aA0wr&sig=fGfpZfr6hah0853yANx4JQ-UhOk&hl=en&sa=X&ei=vbUkVZXfB8_X8gWykIGoDw&redir_esc=y#v=onepage&q=naa%2Fnaca%2045-100%20airfoils&f=false

Apparently the the NACA 45-100 airfoil originated from the NACA 23 series airfoil, and looking at the BL17.5 airfoil(which people have said was apparently NAA's in house designation of the NACA 45-100 airfoil), I think I

agree.

 

Alittle food for thought :

 

 

http://forum.warthunder.com/index.php?/topic/95677-is-the-fw-190-under-performing-in-maneuverability/?p=1765367

Those Lift load calculations you copied and pasted here of the P51D are done with a loadout of 5034Kg / 11 000 pounds, it must have had drop tanks to weigh that much. Only reason I can see someone calculating the P51 with drop tanks would be to purposely skew the numbers in the d9's favour. Odd to say the least.

 

If you take the fighter load out of the P51D - http://www.wwiiaircr...-na-46-130.html - it shows 4359.48Kg / 9 611 pounds.

 

Now if you calculate the lift load using that weight for the P51 you get -

 

Aircraft weight / Wing surface area / Cl max = lift loading

 

4359.48Kg / 21.64 m^2 / 1.28 = 157.38kg/m²

 

 

 

 

 

 

 

http://bbs.hitechcreations.com/smf/index.php?topic=212801.105

As I re-read this thread I noticed several incorrect assumptions:

First the NACA/NAA 45-100 airfoil remained throughout the XP-51, 51A/B/C/D/K and A-36.

The Wing dropped about 3" from the A to the B/C/D/K but the leading edge plan view from ~WS 61 to Root Chord steepened with the D/K wing and also changed LE incidence in that region. The outboard twist remained the same all the way to the tip chord. For the B the washout was -0.5degree at root, 0.75 at ~ WS 50, thence to -0.85 degree at tip. For the D the washout was +0.25 degree at Root, 0.6 degree at ~ WS 61, thence constant to -0.85 degree at tip.

 

The two D CLmax for the NACA/NAA 45-100 was about 1.6 but of course reduced as function of Aspect Ratio. The wind tunnel and flight tests demonstrated a CLmax at approximately 1.4+ for high Reynolds no and about 1.5 for wheels up level flight Stall (Clean - no flaps) While the CLmax with Flaps was close to 1.7 (2-D), the CD of NACA/NAA 45-100 was significantly higher at CLmax. (I don't have my notes in front).

 

I have not seen any publication of any test that suggest a change in Oswald efficiency for any of the P-51 series including the XP-51F/G/J and P-51H with straight leading edge NACA 66-(1.8) 15.5 (at root chord) and 12% at tip.

 

I've read all the informed opinions and have a few thoughts based on research, knowledge of Flight Dynamics factors related to conventional aircraft, and opinions based on the above. First - I have no idea regarding the modeling but I agree just about everything I have read by Dtango and Widewing.

 

The primary distinctive unknown factor between the B and D is how does the difference in Inboard wing incidence as a result of the Leading edge difference in Incidence between the Root and the intersection with the common wing around WS 61? The second question in my mind is what the difference in yaw behavior at low to medium speeds might be between turtledeck B and teardrop canopy D versions?

 

In real life high G turns at corner speeds the application of yaw correction, finite and undefinable, should result in delta trim drag (Ditto and even moreso for a 109 with slats deployed - the local wingtip CL for the high wing is greater than lower wing in greater proportion I would think than for a P-51 which in turn requires slightly more rudder input (and drag) to carve the turn)?

 

After those considerations, the model (IMO) requires derivation with respect to time the following:

Change of drag (dominated by pressure drag as function of lift/AoA) as function of CL as change to velocity and angle of attack occurs from the velocity entering the turn to steady state sustainable velocity in a constant altitude, constant speed turn.

Change to propeller efficiency as the aircraft reduces speed. Shouldn't be much but worth exploring along with trim drag contributions to the T=D equations.

 

As to some of the other stuff commented on here. A Bf 109 (and FW 190) terminal dive speed was less than that of a 51 and a 47. Having said that, with a head start it might get away but the test docs state that 475mph at ~ .72M for 1G dive limit was about all the average 109 should attempt. In addition, the 51 did not suffer the lower speed (than .65 to .72M) compressibility issues that all conventional NACA 23xxx in same 12 to 15% T/C conventional airfoils suffered. The laminar flow wing with T/C at ~ 39% made a huge difference in the drag rise regime.

 

The other complications when modeling dissimilar engines is the plug of HP as a function of altitude and gear stages when comparing say a P-51 and a Bf 109 or FW 190.. as well as stick forces for both roll and high G/high speed turns...It seems the gamers need to input a reasonable (and constant) stick force table for each separate combatant.

 

Edited April 8, 2015 by GrapeJam

[Yo-Yo]

Most the time? There are two figures, 1.01 for the NACA 66(1)-212 and 1.22 for the NACA 23012.

 

Then, use this ridiculous 1.01 CL and compute stall IAS for Mustang....

[Solty]

What is this thread about again? Whose argument can out turn the other person?

Hummingbird says that Fw190D9 should out turn the P-51D.

[[DBS]TH0R]

Having a déjà-vu here...

 

In short, all this comes from different Cl values Hummingbird is arguing about. Hence Yo-Yo's last post.

 

 

Hummingbird says that Fw190D9 should out turn the P-51D.

 

Please read again what SiThSpAwN posted. Unless I am mistaken, this is a class act sarcasm. :)

[Solty]

 

 

 

Please read again what SiThSpAwN posted. Unless I am mistaken, this is a class act sarcasm. :)

 

Lol you are right. I was eating my breakfast and didn't really catch it :3

[Hummingbird]

Then, use this ridiculous 1.01 CL and compute stall IAS for Mustang....

 

No because I don't consider it accurate, hence my previous comment regarding these old 1940's test results and their accuracy. They are basically only accurate relative to each other if done side by side, otherwise they are misleading.

 

Now if only you would provide all these sources you keep talking about then it would makes things a lot easier.

 

What I got on both the NACA 23XXX series and NAA/NACA 45-100, NACA 65 & 66 series all yields an average difference in CLmax of .20 in favor of the NACA 23XXX at ALL speeds that these WW2 fighters were flying at. If you have data to contradict this then post it.

 

The realistic CLmax figures for the Fw190 & P-51 wing are ~1.60 and ~1.40 respectively. Every test at the relevant Reynolds number and speeds say the same thing, ESPECIALLY if field surface roughness is considered as the NACA 23XXX series suffers far less here - and let's face it, these birds weren't flying around with waxed & polished wings ;)

Edited April 8, 2015 by Hummingbird

[GrapeJam]

No because I don't consider it accurate, hence my previous comment regarding these old 1940's test results and their accuracy. They are basically only accurate relative to each other if done side by side, otherwise they are misleading.

Oh but you considered it accurate enough to prove the NACA 23015 massive advantage of 0.2, ok.

 

Oh BTW, the P51A in soviet test only reached 480km/h at SL, the real thing reached 508km/h with 75% power, so using the Soviet P51A as baseline for turn time is a bad idea.

Edited April 8, 2015 by GrapeJam

[Hummingbird]

Oh but you considered it accurate enough to prove the NACA 23015 massive advantage of 0.2, ok.

 

Yes, it's accurate for comparing RELATIVE characteristics! *sigh*

 

The story is the same in the other charts which show higher Clmax values, where the NACA 23XXX usually exhibits a CLmax of ~1.60 and the NAA/NACA 45-100, 66 & 65 series one of ~1.4. AGAIN showing a difference of ~0.2.

 

Is all of this really so hard to understand??

[GrapeJam]

Yes, it's accurate for comparing RELATIVE characteristics! *sigh*

 

The story is the same in the other charts which show higher Clmax values, where the NACA 23XXX usually exhibits a CLmax of ~1.60 and the NAA/NACA 45-100, 66 & 65 series one of ~1.4. AGAIN showing a difference of ~0.2.

 

Is all of this really so hard to understand??

Look at the chart for the Naca 23xxx across the board, 1.60 is seriously is reaching it, more like 1.52-1.55.

 

Coincidentally enough, there's a wing CL "free flow" figure for the D9 on the internet: 1.52.

[Hummingbird]

Look at the chart for the Naca 23xxx across the board, 1.60 is seriously is reaching it, more like 1.52-1.55.

 

Yes, at a Re of 1 mill, move up to the Re of real size aircraft, 6-9 mill, and you'll see the NACA 23XXX get into its own and the laminar flow airfoils (i.e. airfoils with a sharp LE and rearward placed point of max TR) start to suffer in Clmax their main benefit being low drag at low AoAs.

 

 

Coincidentally enough, there's a wing CL "free flow" figure for the D9 on the internet: 1.52.

 

Yes, and one of 1.28 for the Mustang. Wanna use those? :doh:

[GrapeJam]

Yes, and one of 1.28 for the Mustang. Wanna use those? :doh:

 

Yeah, at 5000 kg., the P51D maximum take off weight with full internal fuel is 4500kg, notice something wrong?

 

5000kg is 11000lbs, the P51D's take off weight with 2 x 75 US Gals external tanks.

[Hummingbird]

Yeah, at 5000 kg., the P51D maximum take off weight with full internal fuel is 4500kg, notice something wrong?

 

Weight does not affect the CLmax, so what's your point?

 

Either way I don't understand why you would want to use a figure posted without sources on another forum as reference over actual primary source material.

[GrapeJam]

Weight does not affect the CLmax, so what's your point?

 

Either way I don't understand why you would want to use a figure posted without sources on another forum as reference over actual primary source material.

 

Oh apparently external tanks don't affect clmax as well, ok.

 

And anyway, the FW 190 used the 23015.3 airfoil, and I doubt it's exactly the same as the standard NACA 23xx series airfoil.

[Hummingbird]

Oh apparently external tanks don't affect clmax as well, ok.

 

The guy was using CLmax figures from a report on windtunnel tests with scale models, and the figure for the P-51 (which was for a P-51B model btw) was WITHOUT both tanks and racks.

 

So forget about it already.

[Hummingbird]

And anyway, the FW 190 used the 23015.3 airfoil, and I doubt it's exactly the same as the standard NACA 23xx series airfoil.

 

The last two digits refers to the thickness of the airfoil *sigh* i.e. the Fw190 used an airfoil with a thickness ratio of 15.3% at the root. The first two digits refers to the design, which is a NACA 23XXX series, in other words the design is exactly the same as the NACA 23015 :rolleyes:

 

Anything else?

Edited April 8, 2015 by Hummingbird

[GrapeJam]

Considering that at 192kg/m2, the F4U4 had a stall speed of 96mph IAS, with a 41kg/m2 higher wingloading what do you think the 190's stall speed in IAS will be? Considering that they both used the same airfoil?

 

Or we can alway use the F6F-5's wingloading of 184kg/m2 and stall speed of 96 MPH CAS, but that'll be 49kg/m2 higher wingloading.

Edited April 8, 2015 by GrapeJam