Why was the BF109 so slow compared with the P51?

[drgondog]

How many aircraft used it? Other than Messerschmitt (and Bf109 clones), it seems to be approximately no one. This is usually a sign that an airfoil or airfoil family is not competitive or is poorly characterized. This can happen even with modern airfoils, like the Liebeck airfoils, which have a reputation for very poor off-design characteristics and are not widely used.

I've not been able to find much data on the 2R1and 2R2 series, but that may just be because NASA hasn't gotten around to scanning the relevant reports from the NACA (yes, Messerschmitt used NACA airfoils) reports of the 1920s or very early 1930s, when the airfoil was designed.

After digging into Dave Lednicers "The Incomplete Guide to Airfoil Usage" (The Incomplete Guide to Airfoil Usage), I've found these aircraft that used the 2R1 airfoil. The only US aircraft is the Howard DGA. There was a 2R, a 2R1, and a 2R2 series; none were widely used.

• Root Airfoils
  • 'NACA 2R1 14.2
    • 'Avia CS 199',
    • 'Avia CS 99',
    • 'Avia S 199',
    • 'Avia S 99',
    • 'Hispano HA-1109 Buchan',
    • 'Hispano HA-1110',
    • 'Hispano HA-1112 Buchan',
    • 'Messerschmitt Bf 109B',
    • 'Messerschmitt Bf 109C',
    • 'Messerschmitt Bf 109D Dora',
    • 'Messerschmitt Bf 109E Emil',
    • 'Messerschmitt Bf 109F Fredrich',
    • 'Messerschmitt Bf 109G Gustav',
    • 'Messerschmitt Bf 109K',
    • 'Messerschmitt Me 155B'
• 'NACA 2R1 16'
  • 'Kawasaki Ki-100',
    • 'Kawasaki Ki-61 Hien',
    • 'Kawasaki Ki-88',
    • 'Messerschmitt Me 209V1
• 'NACA 2R1 16.5'
  • 'Kawasaki KAL-2'
• 'NACA 2R1 18.5'
  • 'Messerschmitt Bf 110',
    • 'Messerschmitt Bf 161',
    • 'Messerschmitt Bf 162
• 'NACA 2R1 19'
  • 'Messerschmitt Me 321 Gigant',
    • 'Messerschmitt Me 323 Gigant' ],
• 'NACA 2R2 12' => [
  • 'Howard DGA-11',
    • 'Howard DGA-12',
    • 'Howard DGA-15',
    • 'Howard DGA-8',
    • 'Howard DGA-9'
• Tip Airfoils
• NACA 2R1 10
  Messerschmitt Me 321 Gigant,
  Messerschmitt Me 323 Gigant
  
  NACA 2R1 11
  Messerschmitt Bf 109B,
  Messerschmitt Bf 109C,
  Messerschmitt Bf 109D Dora,
  Messerschmitt Bf 109E Emil,
  Messerschmitt Bf 110,
  Messerschmitt Bf 161,
  Messerschmitt Bf 162
  
  NACA 2R1 11.35
  Avia CS 199,
  Avia CS 99,
  Avia S 199,
  Avia S 99,
  Hispano HA-1109 Buchan,
  Hispano HA-1110,
  Hispano HA-1112 Buchan,
  Messerschmitt Bf 109F Fredrich,
  Messerschmitt Bf 109G Gustav,
  Messerschmitt Bf 109K,
  Messerschmitt Me 155B
  
  NACA 2R12
  CVV 7 Pinocchio
  
  NACA 2R2 12
  Howard DGA-11,
  Howard DGA-12,
  Howard DGA-15,
  Howard DGA-8,
  Howard DGA-9

Dave Lednicer can correct me but IIRC the NACA 2R1 was equivalent to NACA 23xxx airoil with 14.2% thickness. The camber (IIRC) was modified to achieve desirable pitching moment and stall characteristics.

The discussion regarding 'obsolete' airfoils (such as NACA 23xxx and 22xx and 24xx) are not really pertinent and only when compared to the NAA/NACA 45-100 low drag performance do they arise.

What characterized the 45-100 was the relative 'sharper lower slope front half from LE to T/C Max' and the world class attention to detail in the production of the wing. In particular the flush rivets, tight butt joints, flat (Vs wavy) surface from root to tip, surface prep and finish from LE to ~40% chord.

The net result (comparatively)
1. The production tooling and jigs were invested in, designed and built, to produce 'same with high quality' wings.
2. The putty, fill, prime, sand, paint to 40% Chord, top and bottom provided a wing with far less built in boundary layer triggers. Conversely, when the leading edge of top surface of the wing was abused, the resulting early triggers of BL transition from laminar to transitional (But comformal) to full blown turbulent/separated flow was premature.

1. The resulting airflow slipstream characteristics differed from say a NACA 0015 or 23015.5 by the velocity gradient from LE to T/C max. The velocity over the other airfoils with similar thickness - peaked - anywhere from 20-30% of chord, whereas the 45-100 peaked at 40-50% of chord. Additionally the shape of the pressure distribution, while not constant,had a more ordely value of Pressure coefficient as a function of chord.
2. Among the virtues of the different pressure distribution for the 45-100 were"
delayed shock wave formation, further aft, with less movement of CP

• no severe pitch down moment due to shock wave, less loss of pressure distibution aft of shock wave
• Lower velocity gradient to peak value delayed the adverse presure gradient to finally transition from attached, but orderly transition turbulence to full blown and chaotic turbulent region.

IMO, these are reasons why the relatively 'fat' P-51 wing compared to the P-47 and Spitfire had the same Mc ~ 0.75M.

Summary,
Laminar flow vanishes (FOR ALL conventional WWII airfoils) shortly after it begins following LE stagnation point. RN for flat plate transition is ~ 500,000. RN for 100 mph for Mustang is ~ 6x10^7
Transition turbulent flow, still attached to the contours of the wing, although of larger profile compared to the laminar boundary layer continues about 20% more cord than the conventional NACA 23xxx type airfoil. This has the effect of producing a lower profile drag to the free stream.

These comments apply to the region of zero lift through low/medium lift coefficient common to cruise. At high angles of attck the 23xxx will have Lower drag than the 45-100.

 

[Peter Gunn]

Much like a Mustang whose second stage on the blower just kicked in is about where this went over my head...

 

[drgondog]

Well, then, my friend, you need to have another look.

As for the P-51/109 question ... cooling drag. Like the Spitfire, the 109 never actually got Meredith to work despite trying very hard to get it to, therefore binned Meredith and assumed it must've been the wing which, frankly, it wasn't (according to Lee Atwood).

Well, that arguement did not serve Atwood very well, and launched in the great Horkey-Atwood-Schmued hissing contest. Atwood, if honestly and objectively ranking the features - and assigning credit to the thought leaders - would rank Meredith slightly higher than Horkey's wing AND cooling system design plus impovements - then Schmued's passion for low velocity gradient shape, then Kindelberger/Smithson passion for advanced production techniques.

Many questions remain unanswered. Some facts appear consistent relative to Meredith's and Gothert reports concerning cooling drag - namely Atwood was Chief, Engineering from published reports through start of B-25 design. (1934-1939).

The first General Arrangement drawing showing coolant and oil cooler radiators was P-509-1 dated 3-10-40. Scmued managed that drawing, Alger's name is in the title block.
Atwood accompanied the P-509 Drawings, the P-509 Specifications and the P-509 Performance Estimates and Horkey and Rice and certainly aware of the feature and ultimately responsible for signing off the complete package.

The P-509 radiator scheme of imbedded radiators with oil cooler on top and long extended front scoop under wing,under pilot - survived past first detailed specification (1620) into first week of May, when P-509 morphed to 73X with circular matrix and oil cooler incenter and intake scoop located behind flap line.
Those detail changes evolved with wind tunnel testing supervised by Horkey.

Atwood stated in the food fights that he presented Horkey with Meredth and Gothert infomation. Horkey said 'did not'.

Horkey and team both developed the 45-100 wing and the intake, duct plenums, internal vanes of 73x and all successive Mustang designs.

Atwood is correct to state that the Meredith design ultimately conservered 150-200 Hp over a completely open exit flap at top speed at medium altitudes but was not as advantageous above 25-30K due to necessary increases to CL.

Horkey is correct to state that the wing enabled 20+mph better top speed due to drag characteristics at high mach numbers. Equally important was German assessment that the attached flow extended 20% farther than conventional wings (No, Not Laminar Flow, but turbulent transition flow before full blown destruction of the boundray layer). The resultant profile drag was demonstrably less.

What has little light shed upon the subject is the simple fact that X73, as fast as it was, had a flawed Meredith system and lower exhaust gas thrust benefits due to flawed exhaust stack design. Did Atwood take credit for those issues?

The combination of Shenstone, PhD from RAE, wind tunnel testing and changes by Horkey and Allison consultancy cleaned up the X73 to functioning ( and faster) XP-51.. and steadily improved through P-51H

So, at the end of the day didn't Atwood just support the theory of imbedded radiators, intake ducts and exits, cited by Meredith and attempted onvirtually every high performance in-line engine fighter on the horizon? Why was he different from Messerschmitt and Hawker and Supermarine designer attempting to reduce cooling drag?

I submit that the devil is in the details migrating from 'idea' to design results and constant improvements. The actual results belong to Horkey and the Technical Department Chief Waite and new Chief, Engineering Rice. Oh, and Schmued who kept the changes 'clean.

Take what you want, leave the rest.

 

[Daggerr]

"...manifold pressure is also a function of compression ratio,..." To clarify, manifold pressure is a function of supercharger compression ratio (boost) but not of engine compression ratio. That is cylinder pressure (BMEP).

The maximum manifold pressure (boost) that an engine is able to operate on is definitely a function of compression ratio of the piston engine.
The higher the compression ratio of the piston engine, the lower the manifold pressure (boost) that can be used. Greg does not seem to understand that in his video about difference in boost between Merlin and DB engines.

As Tomo Pauk already more or less indicated: Greg has insufficient knowledge of the physics and thermodynamics of a supercharged piston engine.

When Greg talks about boost and power it is usually partly or completely wrong what he says. He usually misses the real point.
Greg probably does not understand what causes pre-ignition and detonation. AFAIR he never talks about it in his videos.
Greg probably does not understand the impact of ambient temperature and manifold temperature on pre-ignition and detonation. AFAIR he never talks about temperature in his videos.
In several of his videos he mentions ata, but admits that he does not really know what it is. Could be 1000 mbar or 1013 mbar or something else, he says. It is actually 1 kgf/cm2, which is 981 mbar.

Greg should read some books about supercharged piston engines. He could start with:
Stanley Hooker: 'The Performance of a Supercharged Aero Engine',
Calum Douglas: 'The Secret Horsepower Race'.

And of course Greg should reduce his extreme bias towards the P-47.

 

[SaparotRob]

While you're all trashing him, a lot of topics he covered would never have come up here. It would have been assumed everyone knew it already. The P-47's inexplicable lack of range (or was it?). the Fighter Mafia, the drop tank saga, super pistons, Material Command and General Doolittle's involvement.
He's this century's Caidin

 

[tomo pauk]

As Tomo Pauk already more or less indicated: Greg has insufficient knowledge of the physics and thermodynamics of a supercharged piston engine.

What is my greatest peeve with Greg: inability to consume/accept any feedback that is not affirmative to his work.

 

[Engineman]

In some respects, discussing the performance of aircraft and their engines demands a pretty high bar of technical knowledge if the true detail is to be accurate.
Myself, I do like to see accurate detail in these matters, but it is rare that it occurs in general discussion. Even well intentioned discussion of technical details is often incorrect in what is deduced, implied or in the quoted detail. A simple example is the term Boost, used in the context of WW2 era piston engines. Boost is often quoted without much explanation, as if it is a common standard but it is not. If anything, Manifold Pressure, or Manifold Air Pressure (MAP) is a more universal term for the pressure in the manifold before the inlet valves and is an absolute pressure, differing only in the unit of measure, normally inches of Mercury for USA equipment and Ata for German. However, Boost is a British term for a relative pressure, relative to standard Sea-level pressure in psi and it should be a positive or negative value. Unfortunately, Boost has become popular slang for all MAP readings and has led to considerable misunderstanding. There is no simple solution to the complication of having different MAP measurements, you have to learn that some are Absolute and some like British Boost are Relative pressure. Experts in the subject will have a good understanding but, if you don't have that you will have to find a table of values.
The important thing is to know what-is-what if you write about it.

Eng

 

[drgondog]

I may have to tread lightly here, there seem to to be a multitude of definitions of "laminar flow" at the time (and even now) in that what can be achieved in the laboratory/wind tunnel is one thing. What can be achieved in the real world, even with a carefully crafted and tended high performance glider with a fiberglass/composite wing, what could be achieved with a metal wing (even with 20 coats of paint sanded between each coat and so on and so on................
There seems to have been something of a misunderstanding in regards to what was needed in surface finish instead of just changing the air foil in order to even get "laminar Flow" to occur over 30-40% of the wings surface. Some wings were referred to as "laminar flow" when they changed the maximum thickness of of the wing from about 30% of cord back to 40-50% of cord. It was a step and some of the best surface finish in the world wasn't going to give much of an improvement in laminar flow if the airfoil wasn't correct but the best airfoil in the world (in the lab/wind tunnel) wasn't going to work in the real world without surface finish standards that were pretty much unobtainable on a service aircraft.

The P-51 did about the best and a 'good' P-51 maintained laminar flow to almost the 40% of cord mark, which is not laminar flow at all according to some people because it didn't maintain it all the way (or even 90%).

The Me 309 or late Ta 152 might have been designed with the intention of getting laminar flow, or at least getting more laminar flow (keeping it non turbulent) over higher percentage of the wing than previous wings but achieving that intention/goal was pretty much beyond the capabilities of the German aircraft industry at the time.

So the question maybe were they designed with the intention/hope of getting laminar flow (or at least a significant reduction in drag) rather than if they achieved it.

I picked this post because there are some good comments to point to the 'why' the NAA/NACAA 45-100 was revolutionary for a Production wing.

First, the progression of stages of airflow needs to be clarified.

Laminar flow. It exists at very low speed for a Mean Aerodynamic Chord of almost 6ft for the Mustang, specifically in the Reynolds Number Range of 600,000, equivalent to slow taxi speed for the P-51. It is characterized by a 'shallow/thin' Boundary Layer attached to the surface of the wing close to the leading edge. Indepepenent of the 'modern airfoils' such as NACA 23xxx or the 45-100, that region near the leading edge is small.

Transitional Turbulent Flow. It exists, exhibits some chaotic internal 'wakes' and disruptions to attached flow,but the flow remains attached with progressively thick boundary layers to 'regions' experience the pressure gradients from the full blown chaotic Turbulent flow state downstream, then

Turbulent Flow with chaotic locations/regions of attached flow but dominantly chaotic with low energy when compared to the free stream passing over it. This 'region' also co-incided with the highest level of math that I took in grad school. Calculus of Variations was a walk in the park compared to attempts to combine Navier Stokes with Chaos theory.

Lednicer, aka aeroweenie please step in.

So, the design of the 45-100 was designed by selecting a perferred pressure distribution, not as high (magnitude) as NACA 23xxx from LE to ~ 30% T/C, but more even through 50-60% before relatively sharp degradation. The region between say 30% and 50% had a better behaved Transitional TURBULENT flow. Not Laminar. Consequently the Center of Pressure for this airfoil was also farther aft so when Mcr was reached near T/C max it was significantly farther downstream than NACA 23xxx. The combination of loss of lift aft of the shockwave and increase of Moment of pitch was mitigated in comparison.

So, that is the 'theoretical' wing. In practice, the wing achieved purty damn gud results to smooth model wing despite camo paint, etc of a production P-51B-1 tested at Langley along with P-63, F4U, F6F, P-40, F4F and P-38. Despite the extraordinary values of the sevice condition vs 'sealed condition' it was a VERY conservative outcome because there was zero Meredith effect to mitigate the unsealed scoop. It effectively added a flat plate inside the cooling duct intake.

So,what?

The wing's extraordinary theoretical design was abedded by the extraordinary manufacturing process to include best practices in tooling, sheet metal fabrication, assy with care to surface imperfections via flush rivets, tight butt joints, etc - resulting in a clean wing at this stage. But the next process of puttying surface gaps, dimples/protrusions - then sanding, priming and painting the leading 40% of the wing ensured consistent results so long as the leading edge was not banged up enough to cause sections of the paint/filler to chip - which would introduce early 'boundary layer trips'.

Now - disclaimer. Certain media experts such as Greq opine that if the Spitfire had the same care in building a wing, that it would be as fast as a Mustang with same power.

I would opine thusly:

If a Spitfire performed elimination of all bumps/protrusions, began with completely covered landing gear, eliminated protruding armaments, etc - collectively would have been be a better design to shave drag.

But with a wing aready 20% thinner, it already had a wing with about the same profile/form drag of the fatter P-51 wing and the friction contributions of the features discussed were not so important as the basic airfoil section.

Greg also kinda forgot about the underwing radiators, the oil cooler, etc when opining about production quality

That said - a liitle more about the Meredith Effect - The REAL advantage, aerodynamically.

I would have humbly beseeched the master internet aerodynamicist, that most of the supriority of the Mustang over Spitfire, Focke Wulf, Messerschmitt, etc was the design and LOCATION of the cooling system. The wing was the icing on the cake, along with an early introduction in fuselage design excellence applying Projective Geometry.


To the 'doubters' of the value of the jet effect of minimum opening exhaust exit scoop, look to the comment in paragraph 3. below.
My calculations yielded about a delta of 180-200 THP loss when exit gate wide open.

Source, detailed flght test performance report, found on P-51B Performance Test

MEMORANDUM REPORT ON
Pursuit Single Engine P-51B-1-NA, AAF No. 43-12093
May 18, 1943

2. High speeds obtained with the oil cooler flap and coolant flap set for automatic operation since there were no provisions on this airplane for selective operation and no time was available for a test installation of a selective control.

True Airspeed M.P.H.​	R.P.M.​	Man. Press. " Hg.​	BHP From Power Chart​	Altitude Ft.​	Coolant Flap Position Inches open From Flush​	Oil Flap Position Inches open From Flush​
(a) Low blower Operation
363​	3,000​	60.5​	1,450​	5,000​	6.0​	W.O (5)​
394​	3,000​	60.5​	1,485​	10,000​	5.0​	3.5​
425​	3,000​	60.5​	1,530​	16,800​	1.5​	1.0​
422​	3,000​	49.0​	1,270​	23,200​	1.0​	Flush​
(b) High blower Operation
422​	3,000​	60.5​	1,270​	23,200​	1.0​	.5​
441​	3,000​	60.5​	1,275​	29,800​	1.0​	Flush​
421​	3,000​	48.0​	985​	35,000​	.5​	Flush​
403​	3,000​	40.7​	815​	38,000​	.5​	Flush​

3. Opening coolant flap wide open from flush position slowed the airplane from 349 M.P.H. I.A.S. to 325 M.P.H. at 18,000 Ft.; opening the oil cooler flap decreased the speed an additional 10 M.P.H. I.A.S.

 

[Deleted member 68059]

I picked this post because there are some good comments to point to the 'why' the NAA/NACAA 45-100 was revolutionary for a Production wing.

First, the progression of stages of airflow needs to be clarified.

Laminar flow. It exists at very low speed for a Mean Aerodynamic Chord of almost 6ft for the Mustang, specifically in the Reynolds Number Range of 600,000, equivalent to slow taxi speed for the P-51. It is characterized by a 'shallow/thin' Boundary Layer attached to the surface of the wing close to the leading edge. Indepepenent of the 'modern airfoils' such as NACA 23xxx or the 45-100, that region near the leading edge is small.

Transitional Turbulent Flow. It exists, exhibits some chaotic internal 'wakes' and disruptions to attached flow,but the flow remains attached with progressively thick boundary layers to 'regions' experience the pressure gradients from the full blown chaotic Turbulent flow state downstream, then

Turbulent Flow with chaotic locations/regions of attached flow but dominantly chaotic with low energy when compared to the free stream passing over it. This 'region' also co-incided with the highest level of math that I took in grad school. Calculus of Variations was a walk in the park compared to attempts to combine Navier Stokes with Chaos theory.

Lednicer, aka aeroweenie please step in.

So, the design of the 45-100 was designed by selecting a perferred pressure distribution, not as high (magnitude) as NACA 23xxx from LE to ~ 30% T/C, but more even through 50-60% before relatively sharp degradation. The region between say 30% and 50% had a better behaved Transitional TURBULENT flow. Not Laminar. Consequently the Center of Pressure for this airfoil was also farther aft so when Mcr was reached near T/C max it was significantly farther downstream than NACA 23xxx. The combination of loss of lift aft of the shockwave and increase of Moment of pitch was mitigated in comparison.

So, that is the 'theoretical' wing. In practice, the wing achieved purty damn gud results to smooth model wing despite camo paint, etc of a production P-51B-1 tested at Langley along with P-63, F4U, F6F, P-40, F4F and P-38. Despite the extraordinary values of the sevice condition vs 'sealed condition' it was a VERY conservative outcome because there was zero Meredith effect to mitigate the unsealed scoop. It effectively added a flat plate inside the cooling duct intake.

So,what?

The wing's extraordinary theoretical design was abedded by the extraordinary manufacturing process to include best practices in tooling, sheet metal fabrication, assy with care to surface imperfections via flush rivets, tight butt joints, etc - resulting in a clean wing at this stage. But the next process of puttying surface gaps, dimples/protrusions - then sanding, priming and painting the leading 40% of the wing ensured consistent results so long as the leading edge was not banged up enough to cause sections of the paint/filler to chip - which would introduce early 'boundary layer trips'.

Now - disclaimer. Certain media experts such as Greq opine that if the Spitfire had the same care in building a wing, that it would be as fast as a Mustang with same power.

I would opine thusly:

If a Spitfire performed elimination of all bumps/protrusions, began with completely covered landing gear, eliminated protruding armaments, etc - collectively would have been be a better design to shave drag.

But with a wing aready 20% thinner, it already had a wing with about the same profile/form drag of the fatter P-51 wing and the friction contributions of the features discussed were not so important as the basic airfoil section.

Greg also kinda forgot about the underwing radiators, the oil cooler, etc when opining about production quality

That said - a liitle more about the Meredith Effect - The REAL advantage, aerodynamically.

I would have humbly beseeched the master internet aerodynamicist, that most of the supriority of the Mustang over Spitfire, Focke Wulf, Messerschmitt, etc was the design and LOCATION of the cooling system. The wing was the icing on the cake, along with an early introduction in fuselage design excellence applying Projective Geometry.


To the 'doubters' of the value of the jet effect of minimum opening exhaust exit scoop, look to the comment in paragraph 3. below.
My calculations yielded about a delta of 180-200 THP loss when exit gate wide open.

Source, detailed flght test performance report, found on P-51B Performance Test

MEMORANDUM REPORT ON
Pursuit Single Engine P-51B-1-NA, AAF No. 43-12093
May 18, 1943

2. High speeds obtained with the oil cooler flap and coolant flap set for automatic operation since there were no provisions on this airplane for selective operation and no time was available for a test installation of a selective control.

True Airspeed M.P.H.​	R.P.M.​	Man. Press. " Hg.​	BHP From Power Chart​	Altitude Ft.​	Coolant Flap Position Inches open From Flush​	Oil Flap Position Inches open From Flush​
(a) Low blower Operation
363​	3,000​	60.5​	1,450​	5,000​	6.0​	W.O (5)​
394​	3,000​	60.5​	1,485​	10,000​	5.0​	3.5​
425​	3,000​	60.5​	1,530​	16,800​	1.5​	1.0​
422​	3,000​	49.0​	1,270​	23,200​	1.0​	Flush​
(b) High blower Operation
422​	3,000​	60.5​	1,270​	23,200​	1.0​	.5​
441​	3,000​	60.5​	1,275​	29,800​	1.0​	Flush​
421​	3,000​	48.0​	985​	35,000​	.5​	Flush​
403​	3,000​	40.7​	815​	38,000​	.5​	Flush​

3. Opening coolant flap wide open from flush position slowed the airplane from 349 M.P.H. I.A.S. to 325 M.P.H. at 18,000 Ft.; opening the oil cooler flap decreased the speed an additional 10 M.P.H. I.A.S.

This is in a file on communication between NAA and the British which was looking at a detailed breakdown of the weight of the P-51 and Spitfire.

It is of course just "a letter", but this was apparently somebodys considered opinion during the war it seems (note carefully
to whom it is addressed...). NB there is a typo in the last bit which is I think that they accidentally picked the Allison
power at its rated altitude, -Probably its F3R V-1710, not F32.- which was a lot lower than 19,000 feet - about 12,500 feet)

 

[drgondog]

This article has plenty of numbers for folks to see. Page 32-33 spit, tempest, 190A-3 and others.

drive.google.com/open?id=1RzSw6SPTOzapjbZza78-fgKdsELrqW-n

That RAE document is one of the best tutorials I have seen regarding Drag discussions.
Tables 6 and 7 raised some questions, however.

One of the major head scratchers for me was the major disparity between Mustang III and Spit IX Thrust values - for the same engine, essentially the same propeller and at the same wind tunnel speed (100fps, 68mph) - of 43pounds Mustang III and 65pounds Spit IX.

Anyone have a notion why? Particularly when the equivalent CDo vs RN is given as Spit IX is (0.0252) vs (0.0175) Mustang III.

Questions to ponder:
If windtunnel data at 100 fps - there is no way to deternmine cooling drag offset due to Meredith design. I also noticed smaller powerlant drag than Spit IX, but only enough to account for cleaner Mustang profile drag for carb and radiator intake

I also noted Mustang X with 'inelegant' radiator scheme had a CDo of 0.0227, much closer to Spit IX , and 0.005 more than P-51B.

 

[drgondog]

This is in a file on communication between NAA and the British which was looking at a detailed breakdown of the weight of the P-51 and Spitfire.

It is of course just "a letter", but this was apparently somebodys considered opinion during the war it seems (note carefully
to whom it is addressed...). NB there is a typo in the last bit which is I think that they accidentally picked the Allison
power at its rated altitude, -Probably its F3R V-1710, not F32.- which was a lot lower than 19,000 feet - about 12,500 feet)

View attachment 745062

Hi Calum i welcome different viewpoints - but suspect the Mustang referenced is the Mustang X. The relative CD0 at 100fps for Spit IX = 0.0252, for Mustang X = 0.0227 and for P-51B = 0.0173.


I just posted the May 1943 P-51B-1 flight test in which there was a drop in TAS from 349IAS to 325IAS at full power/18K

2. High speeds obtained with the oil cooler flap and coolant flap set for automatic opertation since there were no provisions on this airplane for selective operation and no time was available for a test installation of a selective control.

True Airspeed M.P.H.​	R.P.M.​	Man. Press. " Hg.​	BHP From Power Chart​	Altitude Ft.​	Coolant Flap Position Inches open From Flush​	Oil Flap Position Inches open From Flush​
(a) Low blower Operation
363​	3,000​	60.5​	1,450​	5,000​	6.0​	W.O (5)​
394​	3,000​	60.5​	1,485​	10,000​	5.0​	3.5​
425​	3,000​	60.5​	1,530​	16,800​	1.5​	1.0​
422​	3,000​	49.0​	1,270​	23,200​	1.0​	Flush​
(b) High blower Operation
422​	3,000​	60.5​	1,270​	23,200​	1.0​	.5​
441​	3,000​	60.5​	1,275​	29,800​	1.0​	Flush​
421​	3,000​	48.0​	985​	35,000​	.5​	Flush​
403​	3,000​	40.7​	815​	38,000​	.5​	Flush​

Opening coolant flap wide open from flush position slowed the airplane from 349 M.P.H. I.A.S. to 325 M.P.H. at 18,000 Ft.; opening the oil cooler flap decreased the speed an additional 10 M.P.H. I.A.S.

The point that I would make is that only in Flight testing with live engine cooling activity at high speed (above) can one get a 'physical' grasp of the actual magnitude. Low speed wind tunnel results are fine for normal form/profile parasited drag elements but are totally inadequate for exhaust thrust or cooling drag mitigation measurements.

 

[Deleted member 68059]

Hi Calum i welcome different viewpoints - but suspect the Mustang referenced is the Mustang X. The relative CD0 at 100fps for Spit IX = 0.0252, for Mustang X = 0.0227 and for P-51B = 0.0173.


I just posted the May 1943 P-51B-1 flight test in which there was a drop in TAS from 349IAS to 325IAS at full power/18K

2. High speeds obtained with the oil cooler flap and coolant flap set for automatic opertation since there were no provisions on this airplane for selective operation and no time was available for a test installation of a selective control.

True Airspeed M.P.H.​	R.P.M.​	Man. Press. " Hg.​	BHP From Power Chart​	Altitude Ft.​	Coolant Flap Position Inches open From Flush​	Oil Flap Position Inches open From Flush​
(a) Low blower Operation
363​	3,000​	60.5​	1,450​	5,000​	6.0​	W.O (5)​
394​	3,000​	60.5​	1,485​	10,000​	5.0​	3.5​
425​	3,000​	60.5​	1,530​	16,800​	1.5​	1.0​
422​	3,000​	49.0​	1,270​	23,200​	1.0​	Flush​
(b) High blower Operation
422​	3,000​	60.5​	1,270​	23,200​	1.0​	.5​
441​	3,000​	60.5​	1,275​	29,800​	1.0​	Flush​
421​	3,000​	48.0​	985​	35,000​	.5​	Flush​
403​	3,000​	40.7​	815​	38,000​	.5​	Flush​

Opening coolant flap wide open from flush position slowed the airplane from 349 M.P.H. I.A.S. to 325 M.P.H. at 18,000 Ft.; opening the oil cooler flap decreased the speed an additional 10 M.P.H. I.A.S.

The point that I would make is that only in Flight testing with live engine cooling activity at high speed (above) can one get a 'physical' grasp of the actual magnitude. Low speed wind tunnel results are fine for normal form/profile parasited drag elements but are totally inadequate for exhaust thrust or cooling drag mitigation measurements.

Yes its not clear from the letter how the data was arrived at.

 

[ThomasP]

Don't know if this helps, but when I do the calculations (using the methods I use) I get a Cd (not Cd0) of .0175 for the P-51B/Mustang Mk III (clean, w/o racks) taking exhaust thrust and Meredith effects into account, using Vmax of 424/450 mph TAS at 15,000/28,000 ft. This would give a theoretical FPA of ~48.5 ft2 at 100 ft/sec with exhaust thrust and Meredith effect (sort of) taken into account. This would allow for the Cd0/FPA@100 ft/sec values of .0227/63 ft2 for the Mustang Mk X vs .0229/66 ft2 for the Spitfire Mk IX from the same Air Ministry chart dated 9 February 1944.

Maybe . . if I did my math right.

edit sorry, typo - 23,000 ft should be 28,000 ft, fixed it. (Although it works for WEP at 23,000 ft also.)

 

[Peter Gunn]

Man you guys, how many times must I admonish you two ( drgondog D Deleted member 68059 ) about making those hasty generalizations and rather bold statements without data and citations to back them up. Good Grief!!

All kidding aside, either one of you has forgotten more than I know, or will ever know, if you two started a debate in earnest the interwebz might just lawn dart.

 

[drgondog]

Don't know if this helps, but when I do the calculations (using the methods I use) I get a Cd (not Cd0) of .0175 for the P-51B/Mustang Mk III (clean, w/o racks) taking exhaust thrust and Meredith effects into account, using Vmax of 424/450 mph TAS at 15,000/28,000 ft. This would give a theoretical FPA of ~48.5 ft2 at 100 ft/sec with exhaust thrust and Meredith effect (sort of) taken into account. This would allow for the Cd0/FPA@100 ft/sec values of .0227/63 ft2 for the Mustang Mk X vs .0229/66 ft2 for the Spitfire Mk IX from the same Air Ministry chart dated 9 February 1944.

Maybe . . if I did my math right.

edit sorry, typo - 23,000 ft should be 28,000 ft, fixed it. (Although it works for WEP at 23,000 ft also.)

Calum - before we go down a deep rathole, let me understand the assumptions you're dealing with - and let me share some of my immediate thoughts before you respond.

Whether we dispense with CDo vs CDp vs CD, the drag build up of Parasite drag values for the P-51B-1 contained in the Performance Report for Basic Drag are extracted at RN=1.86x10^6. I'll name it Cdp1

That value of Cdp1 is 0.0190.
Correct for new RN
The model was 1/4 scale so the 'L' for MAC is 79.75/4 = 19.93in. Kinematic Vicosity = 0.000333 ft^2 per sec.
The 'V' for RN1 = 1.86x10^6 is for V=120mph; For V=450mph, RN2= 13.15x10^6
Cdp2 = Cdp1*(RN2/RN1)^0.11

Cdp2 = 0.190/1.241 = 0.0153 for incompressible flow

Correct Cdp2 for Mach effect

For 450mph at 28K, M=687mph; 450mph =Vm =450/687 =0.655M
For Mach correction; CDm = Cdp2/(1-Vm^2)^.5
(1-Vm^2}^0.5 =0.755
CDm =.0154/0.755 = .0204 Note. That the multiplier is CD = 1.32 of Incompressible flow drag at 0.655 M

Calculate CL for 450mph
At 8600 pounds (I Guess that is what you are using); CL= W/(0.5*rho*Sw*V^2) at 28000 feet

CL~ 0.176 for V=450 @ 28000;

Calculate Induced Drag
CDi = (CL)^2/pi*AR AR for P-51= 5.89
CDi = (0.176)^2/18.5

CDi = 0.0017

Now CDt = CDm +CDi

CDt = 0.0204 + 0.0017 = 0.0221

This translates to CD/CL of .124 and a HP Required of 1295 for 8600 GW

If it makes sense I will go back and re-do for flight test 441 vs 450 and 8450 vs 8600 once I understand what you were commenting about.

So, additionally - if i calculated correctly Q=208 psf; CDt = D/QS; D=CDt*Q*S = 0.0221*208*233 = 1075 pounds total Drag force.

As you know the above calcs were to find THP Required.

THP Available is quite a bit more complicated because THP losses due to momentum recovery for carb, THP additions via exhaust gas, THP losses due to both propeller efficiency as well as Slipstream drag. If the cooling system exit shutter is opened, there is also cooling drag expressed in mometum recovery calcs. NAA despite claiming some net positive thrust, never used those value in THP avail calcs.

NAA explicitly states 'zero cooling drag' is assumed from SL to Critical Altitude, then 0.0004 at 35000ft, and 0.0010 at 40000ft. These are pressure drag values due to increased angle of attack to maintain level flight.


Also, the NAA stated value for cooling drag Cdp = 0.0064 for climb at all altitudes


What now are we discussing?

 

[Deleted member 68059]

Calum - before we go down a deep rathole, let me understand the assumptions you're dealing with - and let me share some of my immediate thoughts before you respond.

Whether we dispense with CDo vs CDp vs CD, the drag build up of Parasite drag values for the P-51B-1 contained in the Performance Report for Basic Drag are extracted at RN=1.86x10^6. I'll name it Cdp1

That value of Cdp1 is 0.0190.
Correct for new RN
The model was 1/4 scale so the 'L' for MAC is 79.75/4 = 19.93in. Kinematic Vicosity = 0.000333 ft^2 per sec.
The 'V' for RN1 = 1.86x10^6 is for V=120mph; For V=450mph, RN2= 13.15x10^6
Cdp2 = Cdp1*(RN2/RN1)^0.11

Cdp2 = 0.190/1.241 = 0.0153 for incompressible flow

Correct Cdp2 for Mach effect

For 450mph at 28K, M=687mph; 450mph =Vm =450/687 =0.655M
For Mach correction; CDm = Cdp2/(1-Vm^2)^.5
(1-Vm^2}^0.5 =0.755
CDm =.0154/0.755 = .0204 Note. That the multiplier is CD = 1.32 of Incompressible flow drag at 0.655 M

Calculate CL for 450mph
At 8600 pounds (I Guess that is what you are using); CL= W/(0.5*rho*Sw*V^2) at 28000 feet

CL~ 0.176 for V=450 @ 28000;

Calculate Induced Drag
CDi = (CL)^2/pi*AR AR for P-51= 5.89
CDi = (0.176)^2/18.5

CDi = 0.0017

Now CDt = CDm +CDi

CDt = 0.0204 + 0.0017 = 0.0221

This translates to CD/CL of .124 and a Thrust HP Required of 1295 for 8600 GW----> pretty close to the table HP of 1275Hp for 441mph flight test at 29,800ft, GW=8450pounds, 60.5in MP.

If it makes sense I will go back and re-do for flight test 441 vs 450 and 8450 vs 8600 once I understand what you were commenting about.

So, additionally - if i calculated correctly Q=208 psf; CDt = D/QS; D=CDt*Q*S = 0.0221*208*233 = 1075 pounds total Drag force.

As you know the above calcs were to find THP Required.

THP Available is quite a bit more complicated because THP losses due to momentum recovery for carb, THP additions via exhaust gas, THP losses due to both propeller efficiency as well as Slipstream drag. If the cooling system exit shutter is opened, there is also cooling drag expressed in mometum recovery calcs. NAA despite claiming some net positive thrust, never used those value in THP avail calcs.

NAA explicitly states 'zero cooling drag' is assumed from SL to Critical Altitude, then 0.0004 at 35000ft, and 0.0010 at 40000ft. These are pressure drag values due to increased angle of attack to maintain level flight.


Also, the NAA stated value for cooling drag Cdp = 0.0064 for climb at all altitudes


What now are we discussing?

I dont really have a point of any dramatic disagreement, I`d just say the weight of evidence I`ve seen suggests that the P-51 wing was a very significant factor in the lower drag and I`d say its of a significance that is I`d say slightly under-represented by your statement: "most of the supriority of the Mustang... (was the radiator & placement thereof)".

Messerschmitt designed a Laminar flow 109 (which was only not produced because of time and problems with the fact they had such awful finish due to the standard of the labor pool in use by that stage and the effects of dispersal and so on, and captured `51 wings were tested in Germany in full scale tunnels, to great acclaim.

I cant swear on what all these effects add up to, but I feel nervous about saying that the wing was significantly more, or less, of a factor than the radiator was for those reasons.

I cant agree more about the placement of the cooling ducting being as important as the design, but again, I think the radiator was far from perfect too, and probably had significant recirculation at the inlet as it appears to diffuse too fast. I`m sure there were all sorts of compromises involved in that, as NAA must have known that perfectly well too.

I dont want to disagree with you, but I feel even less inclined to disagree with German WW2 aerodyamicists, discussing using American developments to improve their most famous fighter, the stakes for them being wrong in this analysis do not need to be exaggerated, and so I trust them even more than the letter to Camm I already posted. (this is all in English as it was translated by Allied engineers after it was capt... err "borrowed" on permanent loan in 1945.

 

[Engineman]

Great discussion! I think I get the point that true values of the drag or thrust effect of the P-51D cooling duct system are dependent on the system operating with correct mass flows and temperatures. In my mind, the whole assembly is similar to a low ram pressure duct and jet exhaust, with the heat source from the radiator(s).
Am I missing something or, was it not simple to put complete functioning (with hot coolant flows) scoop / duct / radiator(s) / outlet duct assembly in a wind tunnel to measure the actual drag, or thrust and optimise the whole thing?
Cheers

Eng

 

[drgondog]

I dont really have a point of any dramatic disagreement, I`d just say the weight of evidence I`ve seen suggests that the P-51 wing was a very significant factor in the lower drag and I`d say its of a significance that is I`d say slightly under-represented by your statement: "most of the supriority of the Mustang... (was the radiator & placement thereof)".

Messerschmitt designed a Laminar flow 109 (which was only not produced because of time and problems with the fact they had such awful finish due to the standard of the labor pool in use by that stage and the effects of dispersal and so on, and captured `51 wings were tested in Germany in full scale tunnels, to great acclaim.

I cant swear on what all these effects add up to, but I feel nervous about saying that the wing was significantly more, or less, of a factor than the radiator was for those reasons.

I cant agree more about the placement of the cooling ducting being as important as the design, but again, I think the radiator was far from perfect too, and probably had significant recirculation at the inlet as it appears to diffuse too fast. I`m sure there were all sorts of compromises involved in that, as NAA must have known that perfectly well too.

I dont want to disagree with you, but I feel even less inclined to disagree with German WW2 aerodyamicists, discussing using American developments to improve their most famous fighter (this is all in English as it was translated by Allied engineers after it was capt... err "borrowed" on permanent loan in 1945.

View attachment 745194

We are not disagreeing to any significant degree. The reason I cite the cooling system design as more significant (for me) was the the flight test results of opening the exit scoop that I showed above - which basically rendered P-51B speed performance to near that of a P-40 and below a Spit IX.

The NAA reports stating assume CDP=0.0064 for climb conditions in analysis - is closely approximating the CDP = 0.0076 for the complete wing.

Not that a chin radiator selection would result in the entire CDp of 0.0064 but note that with P-51B Base Drag value of 0.0190 adding 1/2 of 0.0064 cooling drag moves it into Mustang X and Spitfire CDP range.

 

[drgondog]

Great discussion! I think I get the point that true values of the drag or thrust effect of the P-51D cooling duct system are dependent on the system operating with correct mass flows and temperatures. In my mind, the whole assembly is similar to a low ram pressure duct and jet exhaust, with the heat source from the radiator(s).
Am I missing something or, was it not simple to put complete functioning (with hot coolant flows) scoop / duct / radiator(s) / outlet duct assembly in a wind tunnel to measure the actual drag, or thrust and optimise the whole thing?
Cheers

Eng

The closest I have seen was the report of the original configuration tested at GALCIT in spring 1940. I doubt that that test was only one with full scale, but the 4-42 test of XP-51B #2 at Ames was with the entire airplane (save wing outboard WS 75) while investigting the Rumble.

After those mods from Ames, the controlling variable was the temperature at the exhaust outlet. Lednicer discusses thoroughly and notes that he assumed 170 degrees F at outlet and also that the inflow BL separation at the top of the plenum with rapid cross sectional increase to radiator(s) face could have been improved.

 

[Engineman]

The closest I have seen was the report of the original configuration tested at GALCIT in spring 1940. I doubt that that test was only one with full scale, but the 4-42 test of XP-51B #2 at Ames was with the entire airplane (save wing outboard WS 75) while investigting the Rumble.

After those mods from Ames, the controlling variable was the temperature at the exhaust outlet. Lednicer discusses thoroughly and notes that he assumed 170 degrees F at outlet and also that the inflow BL separation at the top of the plenum with rapid cross sectional increase to radiator(s) face could have been improved.

Thanks for that. I can see the info in your great book as well. For sure, no-one could expect a perfect design duct system straight-off, but the many stages of progress do seem a bit like "cut it and see". I remain impressed by the final result. I was extremely fortunate to fly in a P-51D two-sticker and it definitely was a valued experience. The airscoop and ducts reminded me of a sharks mouth. What a great design, and what a great engine!

Eng