Yet another P-51 Drag thread

[to_change22]

Hey everyone! I am really sorry to dredge this topic up again. Unfortunately, after reading through some papers on the P-51 as well as all the past forum threads on this topic, I have found myself very confused about the source of the P-51(B or D)'s drag advantages over the Spitfire Mk IX.

First, let's establish some facts.

1. In flight tests with equivalent engines & power-plants, the P-51B/D was 25-30 MPH faster than the Spitfire in level flight.
Drawing from David Lednicer's comparison of WW2 fighters (Source: http://www.wwiiaircraftperformance.org/mustang/Lednicer_Fighter_Aerodynamics.pdf), "with the same version of the Rolls-Royce Merlin engine and propeller installed, the Mustang X was measured to be 23 mph faster than the Spitfire IX....The P-51B is even faster than the Spitfire."
2. The Mustang had lower total drag than the Spitfre. Again, showing a table from Lednicer's article, which uses drag estimates from flight tests and wind tunnels of the era. (Note: he notes in the article that he has been conservatively high with the P-51's drag).

So we know that the Mustang was faster than the Spitfire and had lower total drag.

However, I have been confused out of my mind as to what technical features account for this lower drag. Can this community come to the rescue?

Just one caveat on laminar flow. Quoting Lednicer again: "in service the aircraft was unlikely to have a substantial laminar flow on the wing and transition occurred in the first 15% of the chord". Quoting John Anderson:

So it's unlikely that significantly reduced skin friction drag on the wing accounted for the lower drag unless someone has information that shows that whatever <15% extent laminar flow on the wings was markedly higher than similar extent of laminar flow on the wings of the Spitfire IX.

 

[tomo pauk]

Before people with actual knowledge weight in:
Spitfire's wing was a low-drag affair on itself, mosltly due to the small thickness to chord ratio - 13% at root. P-51's wing (before the 'lightweight Mustangs') was at 16% at root; Spitfire was also with excellent 'resisitance' to the compressibility issues due to the small thickness of the wing. So IMO, just looking at the wing will not provide the answers.
What might give better clues is a look on the stings that result in the 'drag increments', like the weapons' set-up, undercarriage, windscreen, etc. Per the "RAE Technical note No. Aero 1273 (Flight)', a rear view mirror can reduce the speed by as much as 6.75 mph; two cannons protruding can cost 6.25 mph, even a provision for guns' heating going from exhaust stacks cost a bit of speed.
Fixed tailwheel and not fully covered U/C - easy 7-8 mph loss?
Then we have the drag emanating from radiators, where P-51 seems to have a far better solution; I'm not sure what was the speed loss due to this - perhaps another 10 mph?

(drag increments were a thing on the P-51, too, eg. the racks if installed on the P-51A/B/C were 'guilty' for a speed loss of 12 mph at rated altitude)

Once all of this is added up, a loss of 30-40 mph was only to be expected. FWIW, The RAE note says than a Spitfire IX with a throughout clean-up job, but still combat-worthy, would've been able to do 435+ mph.

 

[Deleted member 68059]

Hey everyone! I am really sorry to dredge this topic up again. Unfortunately, after reading through some papers on the P-51 as well as all the past forum threads on this topic, I have found myself very confused about the source of the P-51(B or D)'s drag advantages over the Spitfire Mk IX.

First, let's establish some facts.

1. In flight tests with equivalent engines & power-plants, the P-51B/D was 25-30 MPH faster than the Spitfire in level flight.
Drawing from David Lednicer's comparison of WW2 fighters (Source: http://www.wwiiaircraftperformance.org/mustang/Lednicer_Fighter_Aerodynamics.pdf), "with the same version of the Rolls-Royce Merlin engine and propeller installed, the Mustang X was measured to be 23 mph faster than the Spitfire IX....The P-51B is even faster than the Spitfire."
2. The Mustang had lower total drag than the Spitfre. Again, showing a table from Lednicer's article, which uses drag estimates from flight tests and wind tunnels of the era. (Note: he notes in the article that he has been conservatively high with the P-51's drag).
View attachment 716852
So we know that the Mustang was faster than the Spitfire and had lower total drag.

However, I have been confused out of my mind as to what technical features account for this lower drag. Can this community come to the rescue?

Just one caveat on laminar flow. Quoting Lednicer again: "in service the aircraft was unlikely to have a substantial laminar flow on the wing and transition occurred in the first 15% of the chord". Quoting John Anderson:
View attachment 716853
So it's unlikely that significantly reduced skin friction drag on the wing accounted for the lower drag unless someone has information that shows that whatever <15% extent laminar flow on the wings was markedly higher than similar extent of laminar flow on the wings of the Spitfire IX.

I`m not an aerodynamicist, but, I do "dabble". I have spoken to several VERY senior aerodynamicists about this, and read German wind tunnel tests of the Mustang.

The answer is (as usual) its very complicated and the phrase "it depends" prevails. Which is another way of saying, nobody can provide "an" answer to that.

I am however 95% certain that the following bullet points are fundamentally true.

1) People get very confused about what Laminar flow wings are, are supposed to do in practise, and read things like "usually does not prevail" and then get
very confused and go and confuse others on YouTube.

2) Nobody has a answer for what part of the P-51 gave what percentage gains over a Spitfire IX in drag terms (RR said the Spitfire radiator was 5% of the
total drag of the aircraft, I don't have any other special info to contribute to that)

3) The benefits of "Laminar profiles" is quite a bit more involved than how much of the flow "is" Laminar.

4) The Germans put Mustang wings into wind tunnels, and declared them to be quite astoundingly good, and declared
the drag reduction over the contemporary German wings in use to be from 30 to 40%. They wrote they were very surprised
how far back the laminar flow region extended (25th Jan 1943, by Busemann at TU Braunchschweig)

----
Expanding on the items above:

1a) Laminar flow wings are not expected to have laminar flow over their entire surface, or even most of it, their goal is primarily to have
MORE laminar flow than a standard profile.

1b) The flow regime is totally situationally dependant, there is virtually zero chance that the P-51 wing has significant laminar flow during
aerobatics/dogfight style combat for example.

1c) The P-51 wing is not in fact a standard profile, it was tweaked quite significantly (I think by NAA).

2a) The British didnt know, but did know that the Mustang had a drag force in tests at 100ft/sec of 50lbs instead of 65.5lbs of the Spit IX

3a) There are no reports I have ever seen of brand new replacement Mustangs arriving at squadrons in England and zooming off
into the distance away from their wingmen in well used Mustangs, or for new mustangs suddenly losing a hundred miles of range
after the ground crew forgot to wipe the dead flies off the wings one week.

3b) People who know far more than I do (very senior aerodynamics people who have designed aircraft) think that the P-51 wing
produced a very large portion of its drag reduction through the significant lowering in the size of the frontal pressure resistance area
on the leading edge. (basically because the wing is much pointier at the front than most non-laminar designs). This SHOULD give
you very severe handling problems because once your angle of attack gets high you SHOULD get separation at this sharp(er) leading
edge, the speculation is that the clever tweaking NAA did to the profile used may have somehow kept this issue at a sensible limit.
This is probably the part of the P-51 wing which IS really clever (and almost nobody talks about).

4a) I`ve already posted this elsewhere, but the Germans were desperately copying Laminar flow wings, and a new 109 was designed
with one. So they were very sure it was a big gain. The project was cancelled because the 109`s leading edge slats WERE enough of
a discontinuity (even when retracted) to ruin the effect. Getting around that meant a 30% bigger planform to retain sensible stall
speed without slats, which was not approved due to loss in manufacturing setup as it was a completely different shape entirely
to the standard 109 wing.

So roughly that can be summarised as, "Laminar profiles" have more benefits than having more of the surface in laminar flow than normal
profiles, there is no particularly well documented evidence of really significant unusual degradation of Mustang range or speed in service,
and if they had more time to alter production you`d have seen 109`s with similar wings in combat.

The laminar wing certainly "worked", but for this to be true does NOT mean that it needs 100% laminar flow (which is indeed totally unrealistic
and is totally unobtainable even today).

What IS surprising is that considering the number of flying P-51`s nobody as bloody bothered just taping a load of wool tufts across the
wings and filming the results. (probably Reno people are doing all that kind of stuff but just never tell anyone)

 

[Sisu]

I`m not an aerodynamicist, but, I do "dabble". I have spoken to several VERY senior aerodynamicists about this, and read German wind tunnel tests of the Mustang.

The answer is (as usual) its very complicated and the phrase "it depends" prevails. Which is another way of saying, nobody can provide "an" answer to that.

I am however 95% certain that the following bullet points are fundamentally true.

1) People get very confused about what Laminar flow wings are, are supposed to do in practise, and read things like "usually does not prevail" and then get
very confused and go and confuse others on YouTube.

2) Nobody has a answer for what part of the P-51 gave what percentage gains over a Spitfire IX in drag terms (RR said the Spitfire radiator was 5% of the
total drag of the aircraft, I don't have any other special info to contribute to that)

3) The benefits of "Laminar profiles" is quite a bit more involved than how much of the flow "is" Laminar.

4) The Germans put Mustang wings into wind tunnels, and declared them to be quite astoundingly good, and declared
the drag reduction over the contemporary German wings in use to be from 30 to 40%. They wrote they were very surprised
how far back the laminar flow region extended (25th Jan 1943, by Busemann at TU Braunchschweig)

----
Expanding on the items above:

1a) Laminar flow wings are not expected to have laminar flow over their entire surface, or even most of it, their goal is primarily to have
MORE laminar flow than a standard profile.

1b) The flow regime is totally situationally dependant, there is virtually zero chance that the P-51 wing has significant laminar flow during
aerobatics/dogfight style combat for example.

1c) The P-51 wing is not in fact a standard profile, it was tweaked quite significantly (I think by NAA).

2a) The British didnt know, but did know that the Mustang had a drag force in tests at 100ft/sec of 50lbs instead of 65.5lbs of the Spit IX

3a) There are no reports I have ever seen of brand new replacement Mustangs arriving at squadrons in England and zooming off
into the distance away from their wingmen in well used Mustangs, or for new mustangs suddenly losing a hundred miles of range
after the ground crew forgot to wipe the dead flies off the wings one week.

3b) People who know far more than I do (very senior aerodynamics people who have designed aircraft) think that the P-51 wing
produced a very large portion of its drag reduction through the significant lowering in the size of the frontal pressure resistance area
on the leading edge. (basically because the wing is much pointier at the front than most non-laminar designs). This SHOULD give
you very severe handling problems because once your angle of attack gets high you SHOULD get separation at this sharp(er) leading
edge, the speculation is that the clever tweaking NAA did to the profile used may have somehow kept this issue at a sensible limit.
This is probably the part of the P-51 wing which IS really clever (and almost nobody talks about).

4a) I`ve already posted this elsewhere, but the Germans were desperately copying Laminar flow wings, and a new 109 was designed
with one. So they were very sure it was a big gain. The project was cancelled because the 109`s leading edge slats WERE enough of
a discontinuity (even when retracted) to ruin the effect. Getting around that meant a 30% bigger planform to retain sensible stall
speed without slats, which was not approved due to loss in manufacturing setup as it was a completely different shape entirely
to the standard 109 wing.

So roughly that can be summarised as, "Laminar profiles" have more benefits than having more of the surface in laminar flow than normal
profiles, there is no particularly well documented evidence of really significant unusual degradation of Mustang range or speed in service,
and if they had more time to alter production you`d have seen 109`s with similar wings in combat.

The laminar wing certainly "worked", but for this to be true does NOT mean that it needs 100% laminar flow (which is indeed totally unrealistic
and is totally unobtainable even today).

What IS surprising is that considering the number of flying P-51`s nobody as bloody bothered just taping a load of wool tufts across the
wings and filming the results. (probably Reno people are doing all that kind of stuff but just never tell anyone)

Yes, the Reno racing people have done some extensive tuft testing on a couple of Mustangs over the years. A lot was done and a lot of changes tried to both wing profiles and to wing root fairings and such. As always some stuff worked and some did not -- like combat maneuvering the actual Reno racing environment is impossible to totally replicate at the home airfield. An interesting thread could be put together discussing some of the drag-related discoveries by the Racers over the years.

 

[pbehn]

Sadly for those who love the Spitfire, and I am one, it just isnt as sleek as it looks. If you look from head on you ca see how much more intrusive the Spitfires intakes are, and how much of the wing they cover. That air intake looks aerodynamic but isnt as good as the P-51 set up. There were attempts tp clean up the Spit with the Mk III, but they preferred to have more Mk Vs and then Mk IXs instead of a better Mk VII or VIII.

 

[to_change22]

I`m not an aerodynamicist, but, I do "dabble". I have spoken to several VERY senior aerodynamicists about this, and read German wind tunnel tests of the Mustang.

The answer is (as usual) its very complicated and the phrase "it depends" prevails. Which is another way of saying, nobody can provide "an" answer to that.

I am however 95% certain that the following bullet points are fundamentally true.

1) People get very confused about what Laminar flow wings are, are supposed to do in practise, and read things like "usually does not prevail" and then get
very confused and go and confuse others on YouTube.

2) Nobody has a answer for what part of the P-51 gave what percentage gains over a Spitfire IX in drag terms (RR said the Spitfire radiator was 5% of the
total drag of the aircraft, I don't have any other special info to contribute to that)

3) The benefits of "Laminar profiles" is quite a bit more involved than how much of the flow "is" Laminar.

4) The Germans put Mustang wings into wind tunnels, and declared them to be quite astoundingly good, and declared
the drag reduction over the contemporary German wings in use to be from 30 to 40%. They wrote they were very surprised
how far back the laminar flow region extended (25th Jan 1943, by Busemann at TU Braunchschweig)

----
Expanding on the items above:

1a) Laminar flow wings are not expected to have laminar flow over their entire surface, or even most of it, their goal is primarily to have
MORE laminar flow than a standard profile.

1b) The flow regime is totally situationally dependant, there is virtually zero chance that the P-51 wing has significant laminar flow during
aerobatics/dogfight style combat for example.

1c) The P-51 wing is not in fact a standard profile, it was tweaked quite significantly (I think by NAA).

2a) The British didnt know, but did know that the Mustang had a drag force in tests at 100ft/sec of 50lbs instead of 65.5lbs of the Spit IX

3a) There are no reports I have ever seen of brand new replacement Mustangs arriving at squadrons in England and zooming off
into the distance away from their wingmen in well used Mustangs, or for new mustangs suddenly losing a hundred miles of range
after the ground crew forgot to wipe the dead flies off the wings one week.

3b) People who know far more than I do (very senior aerodynamics people who have designed aircraft) think that the P-51 wing
produced a very large portion of its drag reduction through the significant lowering in the size of the frontal pressure resistance area
on the leading edge. (basically because the wing is much pointier at the front than most non-laminar designs). This SHOULD give
you very severe handling problems because once your angle of attack gets high you SHOULD get separation at this sharp(er) leading
edge, the speculation is that the clever tweaking NAA did to the profile used may have somehow kept this issue at a sensible limit.
This is probably the part of the P-51 wing which IS really clever (and almost nobody talks about).

4a) I`ve already posted this elsewhere, but the Germans were desperately copying Laminar flow wings, and a new 109 was designed
with one. So they were very sure it was a big gain. The project was cancelled because the 109`s leading edge slats WERE enough of
a discontinuity (even when retracted) to ruin the effect. Getting around that meant a 30% bigger planform to retain sensible stall
speed without slats, which was not approved due to loss in manufacturing setup as it was a completely different shape entirely
to the standard 109 wing.

So roughly that can be summarised as, "Laminar profiles" have more benefits than having more of the surface in laminar flow than normal
profiles, there is no particularly well documented evidence of really significant unusual degradation of Mustang range or speed in service,
and if they had more time to alter production you`d have seen 109`s with similar wings in combat.

The laminar wing certainly "worked", but for this to be true does NOT mean that it needs 100% laminar flow (which is indeed totally unrealistic
and is totally unobtainable even today).

What IS surprising is that considering the number of flying P-51`s nobody as bloody bothered just taping a load of wool tufts across the
wings and filming the results. (probably Reno people are doing all that kind of stuff but just never tell anyone)

Hey Calum, thanks for the reply. First off just wanted to say that I'm a huge fan of your book We need more books like it!

I'm going to group your points 1 & 3.

• First off, I'm totally aware that a "laminar flow" airfoils don't need to maintain a laminar boundary layer for 100% of the airfoil section in order to have benefit. Given that the laminar => turbulent boundary layer transition in a "non-laminar" airfoil usually occurs very early in the chord, an airfoil that maintained ~40% laminar flow would be a huge achievement! And according to the CFD models from Lednicer, ~40% of chord is where the transition from laminar => turbulent happens **in theory** (see below for a picture). But Anderson and Lednicer's point is that the wing likely did not maintain laminar flow to anything beyond 15% of chord in level-flight due to all the reasons that have been mentioned in the past (surface roughness being the main one).
  • Now, like I mentioned in my original post, a 15% of chord laminar run might be much higher than other fighters like the Spitfire Mk IX, but I haven't seen any data on where the laminar - turbulent transition happened on the NACA 22xx series that was used for the Spitfire. Do you know if the laminar flow extent on the P-51 B/D in flight was likely *higher* than contemporary fighters (even if it did not get to the 40% that was expected)?
• Re: situation, yes definitely understood. We're talking about level flight conditions only.
• Re: airfoil profile, definitely know that the original B/D airfoils were modified by NAA and are not standard NACA 6-series airfoils. But, the Lednicer paper I cited above has the actual NAA airfoil geometries, so its conclusions hold in my opinion.
• To point 3, could you expand on this? I'm a bit confused by what you said, specifically this quote: "lowering in the size of the frontal pressure resistance area on the leading edge"
  • Here's how I interpreted your point: the pressure distribution of the NACA 6-series profile, and the NAA profile that was based on it, has the maximum degree of suction / negative pressure shifted towards the rear of the airfoil compared to the distribution on a NACA 4-series (eg, like you would see on the Spitfire IX). See below for a picture from Lednicer's paper. This gives you a large region of a favorable (decreasing) pressure gradient which prevents helps encourage laminar flow in the boundary layer. Is that what you meant by "lowering in the size of the frontal pressure resistance area on the leading edge"?
    
    

Regarding Point 4, two questions:

• Do you have a copy of the report? It's cited in Lednicer's paper (see picture), but I have never been able to find it.
  
• Second, the problem I have with that conclusion of the German aerodynamicists is that it's still based on wind-tunnel testing, which we know give misleading results regarding the extent of laminar flow which prevails on the wing unless it's a full-scale wing model and the researchers haven't taken special care to try and make the wing "smoother" than it would be in service. I'd need to actually read the report to understand the conditions under which the drag values were found.

Anyways thanks for the detailed response!

 

[PAT303]

The Spit didn't have any one thing attributing to drag but a lot of small things added up. The cannons, the outboard cannon stubs, the cannons blisters, the angle of the windscreen, the rear view mirror, the uncovered undercarraige, the tail wheel and oil cooler, each was a small amount but together were worth about 25mph, the radiators don't seem to get much attention in all the reports I've read which is surprising given how much attention they get. A good example of the difference small details make is when you compare the Seafire MkII to the Seafire MkIII, the MkIII has the cannon stubs removed, ejector exhausts and MkV Hispano's in slimline blisters.

 

[Deleted member 68059]

Hey Calum, thanks for the reply. First off just wanted to say that I'm a huge fan of your book We need more books like it!

I'm going to group your points 1 & 3.

• First off, I'm totally aware that a "laminar flow" airfoils don't need to maintain a laminar boundary layer for 100% of the airfoil section in order to have benefit. Given that the laminar => turbulent boundary layer transition in a "non-laminar" airfoil usually occurs very early in the chord, an airfoil that maintained ~40% laminar flow would be a huge achievement! And according to the CFD models from Lednicer, ~40% of chord is where the transition from laminar => turbulent happens **in theory** (see below for a picture). But Anderson and Lednicer's point is that the wing likely did not maintain laminar flow to anything beyond 15% of chord in level-flight due to all the reasons that have been mentioned in the past (surface roughness being the main one).
  • Now, like I mentioned in my original post, a 15% of chord laminar run might be much higher than other fighters like the Spitfire Mk IX, but I haven't seen any data on where the laminar - turbulent transition happened on the NACA 22xx series that was used for the Spitfire. Do you know if the laminar flow extent on the P-51 B/D in flight was likely *higher* than contemporary fighters (even if it did not get to the 40% that was expected)?
• Re: situation, yes definitely understood. We're talking about level flight conditions only.
• Re: airfoil profile, definitely know that the original B/D airfoils were modified by NAA and are not standard NACA 6-series airfoils. But, the Lednicer paper I cited above has the actual NAA airfoil geometries, so its conclusions hold in my opinion.
• To point 3, could you expand on this? I'm a bit confused by what you said, specifically this quote: "lowering in the size of the frontal pressure resistance area on the leading edge"
  • Here's how I interpreted your point: the pressure distribution of the NACA 6-series profile, and the NAA profile that was based on it, has the maximum degree of suction / negative pressure shifted towards the rear of the airfoil compared to the distribution on a NACA 4-series (eg, like you would see on the Spitfire IX). See below for a picture from Lednicer's paper. This gives you a large region of a favorable (decreasing) pressure gradient which prevents helps encourage laminar flow in the boundary layer. Is that what you meant by "lowering in the size of the frontal pressure resistance area on the leading edge"?
    View attachment 717004

Regarding Point 4, two questions:

• Do you have a copy of the report? It's cited in Lednicer's paper (see picture), but I have never been able to find it. View attachment 717005
• Second, the problem I have with that conclusion of the German aerodynamicists is that it's still based on wind-tunnel testing, which we know give misleading results regarding the extent of laminar flow which prevails on the wing unless it's a full-scale wing model and the researchers haven't taken special care to try and make the wing "smoother" than it would be in service. I'd need to actually read the report to understand the conditions under which the drag values were found.

Anyways thanks for the detailed response!

Yes I`ve got that Busemann report and one other in addition, but they are all in German (soemtimes such reports came with translations done by Allied people in 45/46), and they are also pretty academic. So quite painful to work through, but here is the concluding paragraph and here google translate will help you.

There is of course another argument here when the discussion gets to this type of stuff, which is how turbluent was THIS windtunnel ? Early windtunnels (1920`s especially)
were very turblent and this "masked" many aerodynamic benefits (Shenstone considers this was a major reason why Biplanes survived so long). At that point
its more or less time for me to sign out of the conversation as I dont posess the training to understand how that may influence this type of wing test.

 

[drgondog]

First - a hat off for Calum as he pulled together the underlying science based on sources and 'expert witnesses'. Another one to Tomo who pointed out significant external feaures which disrupted harmony of 'tranquil flow' of the Spifire vs P-51B. Always a hat off to Lednicer for his very clear analysis using CFD combined with very sophisticated Navier Stokes mesh. David lurks this site and always offers concise opinions whenhe weighs in.

The Wing, the second order conical design of the fuselage ahead of the lifting line and the Meridith based cooling system are the primary aerodynamic drag reduction features. The exhaust gas thrust which increased with MP was a very important Thrust feature (presuambly near same with same Mark engine and Boost with Spitfire - but higher than 109G for example).

Lednicer also pointed out the Spitfire IX windshield as an unusually large contributor to Sit drag and increasing important in trans Mach region. The P-51B was also suspect in this area as both exhibited stagnation flow at the base of the windshield. In one area I depart from Lednicer in viewpoint, namely Total Drag of P-51D higher than P-51B. From both the wind tunnel based comparisons of NAA analysis of P-51B and D and H, the Drag build up of parasite drag for the fuselage/cockpit combination decreased slightly from B to D. For the empennage, the B/D remained same. The wing between B/D same save the LE change from WS 61 inboard to fusleage - more wetted area for D and local change of relative AoA. The P-51H wing was different - not only larger in area but also in airfoil - to another High Speed/Low drag NACA 66 series with lower parasite drag despite increase in wing area.

Why do I disagree on the Lednicer B/D discussion? First the drag build ups as contained in the NAA Performance Analysis reports for P-51B-1 and P-51D-5 when compared at same RN favor the P-51D.

Second, the TAS of the P-51D-15 with racks at 9700+ pounds (no fuse tank fuel) at 67"MP

P 51D Performance Test

and P-51B-15 with racks at 9300 pounds - full combat weight at 67"MP

P-51B Performance Test

Note that top speed vs altitude of the B @75"MP without racks is ~12mph faster than with racks when P-51B-15 GW = 9700+.

http://www.wwiiaircraftperformance.org/mustang/p-51b-24771-level-blue.jpg

When you evaluate a.) Gross weight at take off, b.) Same engine/same MP/same altitude, c.) external differences (wing rack delta) - two facts stand out.

At approximately the same gross weight at takeoff to compare both models with approximately the same Induced drag (actually giving the B an advantage with 300+ pound lighter configuration) the D was 16mph faster (444 vs 428) at 67"MP and ~ same critical altitude cited.

We have no available 75" test for the P-51D with or without racks but are pretty sure the speed advantage oy th ETO permitted 72"MP takes the P-51D with drained 8 gal tank, but full wing tanks (9700 vs 10200 pounds at take off) will be in 450mph at 24K at 72" with racks - over Berlin or Munich or Prague.

Now for Spit IX at 18" Boost. Note - these are the best speed runs (that I could find) with external features covered to reduce drag, and specifically no mention of C/L fuel tank rack.

Spitfire LF Mk IX Speed Trials Note 411mph at FTH of 21000 feet. The P-51D-15 at same 95% Max Internal gross take off weight has top speed at 20,000 feet (6ooo below FTH) of 421 With both external racks and increases speed advantage with altiutude for same boost - this report

Spitfire H.F. IX Trials

Shows that best gear ratio of 0.477:1 achieves 413mph at same boost, RPM, fraction of GW and critical altitude versus 442mph of P-51D-15 WITH wing racks - nearly a 40mph difference at same altitude, but clean Spit versus wing rack handicapped P-51D-15.

Summary of my personal conclusions:
1. The NAA/NACA 45-100 High Speed/Low Drag wing compared to Spitfire NACA 2213 wing, exhibits a distinct 'bucket' in profile drag - below the Spit in the low CL range for high speed flight. The effective profile drag of the Mustang 16% wing is the same as the Spitfire 13% wing at high speed.
2. The Spit wing despite 3% lower T/C barely edges out the Mustang wing at Vmach>.6 in CDtotal Vs Vmach. Less than 1%.
3. The comparison of frontal view displays the fundamental difference in 'objects' seen by the airflow. The abrupt change in x-section at the nose from spinner to wing root, b.) size and locations the carb and radiator scoops. Additionally the Mustang intake scoop is well aft of lifting line for wing; c.) the bulges in wing and 20mm guns versus 'immersed' 50 cal guns and the 'gap' of main gear wheel well; d.) the 'flat plate/stagnation effect of Spitfire windscreen vs steeply sloped P-51D.
4. Subtle, but important is that all the multiple fuselage 'boo boo's are immersed in prop votex and subject to pressure drag delta higher than P-51 as a result.
5. Last, but extremely important is that the effect of parasite and pressure drag reduction of the Mustang's frontal area, combined with Meridith effect design, approaches zero HP loss to overcome cooling drag.

The last point is subject to debate. Based on my research of NAA docs, reports and memos, the NA-73through 99 'had some' cooling drag at top speed level flight; The P-51B/D 'assumed zero' and so stated based on wind tunnel data; The P-51H achieved slight net thrust at 90" and high(er) exit temperatures at aft scoop compared to B/D.

Final notes; NAA pioneered the application of Projective Geometry as applied to lofting lines development for aircraft. This is the foundation of 2nd degree conical sections mumbo jumbo - but ESSENTIAL to develop an airframe with smoothly increasing, but low, pressure gradient from nose through the lifting line of the wing. Roy Liming and R.K. Weebe devoped the 'bible' and Edgar Schmued was the 'preacher'.

Extending the conversation, the P-51 wing design was founded on the application of Complex Variables and transformation of developed pressure distribution around a rotating sphere - then tranforming back to desired pressure distribution along wing coordinates in real world.

The resulting shape was different from NACA 45-125 20% T/C Laminar Flow wing.
1. As mentioned above the LE radius was smaller than conventional airfoils like the very good NACA 230xx.
2. The max T/C was at 37.5%, approximately 5-10% behind the Spit, Bf 109, P-38, FW 190, F6F, etc wings.
3. Lower velocity gradient from LE to T/Cmax which had the effect of delaying Mcrit and also minimize CP movement during the transition.

 

[Daggerr]

Very interesting topic.

FWIW: I once read (don't remember where) that the low P-51 drag was partly because it happened to be better area-ruled than the Spitfire and others.
That was likely not deliberate, as the area rule was not discovered until 1944 in Germany when it was patented by Otto Frenzl as the 'Querschnittsflächenregel', while in the US it took until 1952 before it was rediscovered by Richard Whitcomb.

I have never understood the point of wind tunnel testing planes or wings or drop tanks or ..... (whatever) at speeds of only 100 ft/s. At such such speed area ruling is unlikely to have a noticeable effect. OTOH I would expect that the benefit of a laminar flow wing would be overestimated, as at low air speed the transition between laminar and turbulent flow is likely to occur further aft of the leading edge.

 

[drgondog]

Very interesting topic.

FWIW: I once read (don't remember where) that the low P-51 drag was partly because it happened to be better area-ruled than the Spitfire and others.
That was likely not deliberate, as the area rule was not discovered until 1944 in Germany when it was patented by Otto Frenzl as the 'Querschnittsflächenregel', while in the US it took until 1952 before it was rediscovered by Richard Whitcomb.

I have never understood the point of wind tunnel testing planes or wings or drop tanks or ..... (whatever) at speeds of only 100 ft/s. At such such speed area ruling is unlikely to have a noticeable effect. OTOH I would expect that the benefit of a laminar flow wing would be overestimated, as at low air speed the transition between laminar and turbulent flow is likely to occur further aft of the leading edge.

Wind tunnel testing was primarily for four purposes.

First, to establish flight related behavior with respect to flow properties and various Cm,Cd and Cl for 'real' simulation.

Second was to build and test improvements to the airframe/wing coniderations prior to fabricatin in real world.

Third, was to define the relationships between CD and R.N.

Fourth, define parasite drag for components of the airframe for later use to predict performance.

100mph was a common benchmark for Brit discussions of drag forces, but Cd as f(R.N.) more common for American counterparts. The latter approach is necessary as most wind tunnel approaches were with scale models - for which a RN based on Mean Aero Chord of a 1/4 model or 1/3 scale model is already decididly less than for full scale in same wind tunnel speed.

 

[drgondog]

Very interesting topic.

FWIW: I once read (don't remember where) that the low P-51 drag was partly because it happened to be better area-ruled than the Spitfire and others.
That was likely not deliberate, as the area rule was not discovered until 1944 in Germany when it was patented by Otto Frenzl as the 'Querschnittsflächenregel', while in the US it took until 1952 before it was rediscovered by Richard Whitcomb.

First applied to F-102 migration to F-106 (IIRC) but area rule 'awareness' existed earlier than 1952

See my discussion above for Descriptive Geometry application to Mustang - and later to F-86 and F-100. Although Area Rule engineering was considered at NAA for F-100, the program was too far advanced to apply.

 

[to_change22]

First - a hat off for Calum as he pulled together the underlying science based on sources and 'expert witnesses'. Another one to Tomo who pointed out significant external feaures which disrupted harmony of 'tranquil flow' of the Spifire vs P-51B. Always a hat off to Lednicer for his very clear analysis using CFD combined with very sophisticated Navier Stokes mesh. David lurks this site and always offers concise opinions whenhe weighs in.

The Wing, the second order conical design of the fuselage ahead of the lifting line and the Meridith based cooling system are the primary aerodynamic drag reduction features. The exhaust gas thrust which increased with MP was a very important Thrust feature (presuambly near same with same Mark engine and Boost with Spitfire - but higher than 109G for example).

Lednicer also pointed out the Spitfire IX windshield as an unusually large contributor to Sit drag and increasing important in trans Mach region. The P-51B was also suspect in this area as both exhibited stagnation flow at the base of the windshield. In one area I depart from Lednicer in viewpoint, namely Total Drag of P-51D higher than P-51B. From both the wind tunnel based comparisons of NAA analysis of P-51B and D and H, the Drag build up of parasite drag for the fuselage/cockpit combination decreased slightly from B to D. For the empennage, the B/D remained same. The wing between B/D same save the LE change from WS 61 inboard to fusleage - more wetted area for D and local change of relative AoA. The P-51H wing was different - not only larger in area but also in airfoil - to another High Speed/Low drag NACA 66 series with lower parasite drag despite increase in wing area.

Why do I disagree on the Lednicer B/D discussion? First the drag build ups as contained in the NAA Performance Analysis reports for P-51B-1 and P-51D-5 when compared at same RN favor the P-51D.

Second, the TAS of the P-51D-15 with racks at 9700+ pounds (no fuse tank fuel) at 67"MP

P 51D Performance Test

and P-51B-15 with racks at 9300 pounds - full combat weight at 67"MP

P-51B Performance Test

Note that top speed vs altitude of the B @75"MP without racks is ~12mph faster than with racks when P-51B-15 GW = 9700+.

http://www.wwiiaircraftperformance.org/mustang/p-51b-24771-level-blue.jpg

When you evaluate a.) Gross weight at take off, b.) Same engine/same MP/same altitude, c.) external differences (wing rack delta) - two facts stand out.

At approximately the same gross weight at takeoff to compare both models with approximately the same Induced drag (actually giving the B an advantage with 300+ pound lighter configuration) the D was 16mph faster (444 vs 428) at 67"MP and ~ same critical altitude cited.

We have no available 75" test for the P-51D with or without racks but are pretty sure the speed advantage oy th ETO permitted 72"MP takes the P-51D with drained 8 gal tank, but full wing tanks (9700 vs 10200 pounds at take off) will be in 450mph at 24K at 72" with racks - over Berlin or Munich or Prague.

Now for Spit IX at 18" Boost. Note - these are the best speed runs (that I could find) with external features covered to reduce drag, and specifically no mention of C/L fuel tank rack.

Spitfire LF Mk IX Speed Trials Note 411mph at FTH of 21000 feet. The P-51D-15 at same 95% Max Internal gross take off weight has top speed at 20,000 feet (6ooo below FTH) of 421 With both external racks and increases speed advantage with altiutude for same boost - this report

Spitfire H.F. IX Trials

Shows that best gear ratio of 0.477:1 achieves 413mph at same boost, RPM, fraction of GW and critical altitude versus 442mph of P-51D-15 WITH wing racks - nearly a 40mph difference at same altitude, but clean Spit versus wing rack handicapped P-51D-15.

Summary of my personal conclusions:
1. The NAA/NACA 45-100 High Speed/Low Drag wing compared to Spitfire NACA 2213 wing, exhibits a distinct 'bucket' in profile drag - below the Spit in the low CL range for high speed flight. The effective profile drag of the Mustang 16% wing is the same as the Spitfire 13% wing at high speed.
2. The Spit wing despite 3% lower T/C barely edges out the Mustang wing at Vmach>.6 in CDtotal Vs Vmach. Less than 1%.
3. The comparison of frontal view displays the fundamental difference in 'objects' seen by the airflow. The abrupt change in x-section at the nose from spinner to wing root, b.) size and locations the carb and radiator scoops. Additionally the Mustang intake scoop is well aft of lifting line for wing; c.) the bulges in wing and 20mm guns versus 'immersed' 50 cal guns and the 'gap' of main gear wheel well; d.) the 'flat plate/stagnation effect of Spitfire windscreen vs steeply sloped P-51D.
4. Subtle, but important is that all the multiple fuselage 'boo boo's are immersed in prop votex and subject to pressure drag delta higher than P-51 as a result.
5. Last, but extremely important is that the effect of parasite and pressure drag reduction of the Mustang's frontal area, combined with Meridith effect design, approaches zero HP loss to overcome cooling drag.

The last point is subject to debate. Based on my research of NAA docs, reports and memos, the NA-73through 99 'had some' cooling drag at top speed level flight; The P-51B/D 'assumed zero' and so stated based on wind tunnel data; The P-51H achieved slight net thrust at 90" and high(er) exit temperatures at aft scoop compared to B/D.

Final notes; NAA pioneered the application of Projective Geometry as applied to lofting lines development for aircraft. This is the foundation of 2nd degree conical sections mumbo jumbo - but ESSENTIAL to develop an airframe with smoothly increasing, but low, pressure gradient from nose through the lifting line of the wing. Roy Liming and R.K. Weebe devoped the 'bible' and Edgar Schmued was the 'preacher'.

Extending the conversation, the P-51 wing design was founded on the application of Complex Variables and transformation of developed pressure distribution around a rotating sphere - then tranforming back to desired pressure distribution along wing coordinates in real world.

The resulting shape was different from NACA 45-125 20% T/C Laminar Flow wing.
1. As mentioned above the LE radius was smaller than conventional airfoils like the very good NACA 230xx.
2. The max T/C was at 37.5%, approximately 5-10% behind the Spit, Bf 109, P-38, FW 190, F6F, etc wings.
3. Lower velocity gradient from LE to T/Cmax which had the effect of delaying Mcrit and also minimize CP movement during the transition.

Hi drgondog , I was hoping you'd pop into the chat.

Let's take this one by one.

The wing.

You state: " The NAA/NACA 45-100 High Speed/Low Drag wing compared to Spitfire NACA 2213 wing, exhibits a distinct 'bucket' in profile drag - below the Spit in the low CL range for high speed flight.".

1. Why does the wing generate this low profile drag? Typically, the low drag "bucket" at low CLs for the NACA 6-series is attributable to the development of laminar flow over a large portion of the chord and the resulting reduction in skin friction drag on the wing. But as I said in my post, both Lednicer and Anderson do not think that substantial laminar flow developed on the P-51 wing in flight. So if not laminar flow, what else accounts for the low profile drag of the wing?

2. Second, do you have a source for this? And if yes, was it a wind-tunnel test or a flight test?

3. You state, "the Spit wing despite 3% lower T/C barely edges out the Mustang wing at Vmach>.6 in CDtotal Vs Vmach. Less than 1%." I didn't understand this, could you rephrase? What is Vmach? Is it the critical mach number?

The radiator

So I want to distinguish between two values: gross cooling drag (assuming the radiator installation produced no thrust) and net cooling drag (net of the thrust produced by the Meredith effect).

1. First, just with respect to gross cooling drag (so let's ignore the Meredith effect), how did the P-51's radiator installation compare to the Spitfire Mk IX? Ie, when you say "zero cooling drag" do you mean gross or net?

2. Second, I've seen conflicting claims on whether the the Meredith effect accounted for any significant thrust. What's the best source on whether it in fact did produce thrust and if so, how much?

The Fuselage

1. Can you expand on the use of conical sections or provide a link to learn more? Why do they result in low drag?

Thanks so much!

 

[to_change22]

Yes I`ve got that Busemann report and one other in addition, but they are all in German (soemtimes such reports came with translations done by Allied people in 45/46), and they are also pretty academic. So quite painful to work through, but here is the concluding paragraph and here google translate will help you.

There is of course another argument here when the discussion gets to this type of stuff, which is how turbluent was THIS windtunnel ? Early windtunnels (1920`s especially)
were very turblent and this "masked" many aerodynamic benefits (Shenstone considers this was a major reason why Biplanes survived so long). At that point
its more or less time for me to sign out of the conversation as I dont posess the training to understand how that may influence this type of wing test.

View attachment 717019

Thanks Calum, here's the translation for those interested.

"In the "Mustang P-51" profile, the transition point has shifted to about x/l = 0.5 to 0.6 in the Re number range up to 4.10^6. This results in savings of 30 to 40% in profile resistance at the profile resistance compared to the iron files NACA 23812 and NACA 4415 with values Ca = 0.08 to 0.20. The favorable Ca range is probably even larger. The curve Ca(a) is normally linear, the curve Cm(Ca) is piecewise linear; the large irregularities otherwise observed in laminar profiles do not occur."

This tells me that the Germans were really excited about the prospects for laminar boundary layers - they noted that the laminar-to-turbulent transition happened at 0.5 - 0.6 of chord - higher than what was measured in NACA tests. But again, I think this runs up against the criticism that both Lednicer and Anderson note. They say that laminar - to - turbulent transition happens in flight < 0.15 of chord. At that point like you said, we need Lednicer himself to comment - someone with more experience in wind tunnels to understand whether this test is representative of flight conditions.

Hence why I was interested in your point #3, which was about benefits of the P-51 airfoil that don't have to do with less skin friction drag due to a laminar boundary layer. So I'll re-up the question:

My question - To point 3, could you expand on this? I'm a bit confused by what you said, specifically this quote: "lowering in the size of the frontal pressure resistance area on the leading edge"

• Here's how I interpreted your point: the pressure distribution of the NACA 6-series profile, and the NAA profile that was based on it, has the maximum degree of suction / negative pressure shifted towards the rear of the airfoil compared to the distribution on a NACA 4-series (eg, like you would see on the Spitfire IX). See below for a picture from Lednicer's paper. This gives you a large region of a favorable (decreasing) pressure gradient which prevents helps encourage laminar flow in the boundary layer. Is that what you meant by "lowering in the size of the frontal pressure resistance area on the leading edge"?

 

[Deleted member 68059]

Thanks Calum, here's the translation for those interested.
yep
"In the "Mustang P-51" profile, the transition point has shifted to about x/l = 0.5 to 0.6 in the Re number range up to 4.10^6. This results in savings of 30 to 40% in profile resistance at the profile resistance compared to the iron files NACA 23812 and NACA 4415 with values Ca = 0.08 to 0.20. The favorable Ca range is probably even larger. The curve Ca(a) is normally linear, the curve Cm(Ca) is piecewise linear; the large irregularities otherwise observed in laminar profiles do not occur."

This tells me that the Germans were really excited about the prospects for laminar boundary layers - they noted that the laminar-to-turbulent transition happened at 0.5 - 0.6 of chord - higher than what was measured in NACA tests. But again, I think this runs up against the criticism that both Lednicer and Anderson note. They say that laminar - to - turbulent transition happens in flight < 0.15 of chord. At that point like you said, we need Lednicer himself to comment - someone with more experience in wind tunnels to understand whether this test is representative of flight conditions.

Hence why I was interested in your point #3, which was about benefits of the P-51 airfoil that don't have to do with less skin friction drag due to a laminar boundary layer. So I'll re-up the question:

My question - To point 3, could you expand on this? I'm a bit confused by what you said, specifically this quote: "lowering in the size of the frontal pressure resistance area on the leading edge"

• Here's how I interpreted your point: the pressure distribution of the NACA 6-series profile, and the NAA profile that was based on it, has the maximum degree of suction / negative pressure shifted towards the rear of the airfoil compared to the distribution on a NACA 4-series (eg, like you would see on the Spitfire IX). See below for a picture from Lednicer's paper. This gives you a large region of a favorable (decreasing) pressure gradient which prevents helps encourage laminar flow in the boundary layer. Is that what you meant by "lowering in the size of the frontal pressure resistance area on the leading edge"?

yep

 

[PAT303]

Another bloody good thread this one.

 

[drgondog]

Hi drgondog , I was hoping you'd pop into the chat.

Let's take this one by one.

The wing.

You state: " The NAA/NACA 45-100 High Speed/Low Drag wing compared to Spitfire NACA 2213 wing, exhibits a distinct 'bucket' in profile drag - below the Spit in the low CL range for high speed flight.".

1. Why does the wing generate this low profile drag? Typically, the low drag "bucket" at low CLs for the NACA 6-series is attributable to the development of laminar flow over a large portion of the chord and the resulting reduction in skin friction drag on the wing. But as I said in my post, both Lednicer and Anderson do not think that substantial laminar flow developed on the P-51 wing in flight. So if not laminar flow, what else accounts for the low profile drag of the wing?

Ed Horkey explained this to me many years ago when I was an aspiring Aero, then followed up in his book "The Real Stuff - The Story Behind the P-51 Mustang" pgs 97-101. That said, he illustrated the difference in pressure distribution, T/C location for boundary layer 'laminar to turbulent' transition point between the NACA 0012 airfoil and the NACA 65.2-016 airfoil - not the similar, but not same NAA/NACA 450-100.

Plate XIV-1 is particularly interesting because it Doesn't use the XP51F/G or P-51H specific NACA 66,2 series. That said, the llustrated transiton(s) displayed two remarkable features: 1.) the 0012 airfoil presented the transition of sharp rise of Cp at LE to ~ -0.42 at ~ 0.2 T/C with gradual reduction in pressure coefficient aft of that point to near zero at 0.8 TC, down to Cp=+0.6 at TE, and 2.) a less steep pressure distribution from LE to peak of Cp=-0.5 at 50% then a sharper drop in pressure to Cp= 0.25 at TE.

PlateXIV-3 displays the associated CD vs CL curves to illustrate the dramatic 'bucket' from CL= -0.5 to CL=0.4 for the NACA 65 vs 0012 wing section.

He further explains in great detail regading the application of Drag Rake testing to explore wake drag to derive pressure drag as f(T/C) wing station showed conclusively that 'laminar flow', if defined as great reduction in the pressure drag (vs friction), then both the NACA 65/66 series and the 45-100 did achieve that state -

Look also to Journal, American Avation Historical Society, Fall 1996 for his "The P-51 The Real Story" pgs 178-189 for is discussions of Wing vs Meridith Effect. It is an excellent article, encompasses what he said in in book and futher explains how the concave x-section of wing aft of max T/C for conventional airfoil realizes negative pressures -"suction" (create incremental drag vector parallel to chord line) normal to the surface creates 'drag', whereas the cusped setion of the NAA airfoils realized POSITIVE pressures on top and bottom wing structure - aft of Max T/C - to achive incremental thrust vectors parallel to chord line. Thus another explanation for reduction of pressure drag.

2. Second, do you have a source for this? And if yes, was it a wind-tunnel test or a flight test?

NAA performed a series of Drag Wake testing to compliment those conducted by NACA in 1942 on XP-51 41-038 and later 41-039. See NACA Advance Confidential Report ACR No. 2K01 for first trenche dated November 1942, See the follow up,more extensive Rake testing for 039 in NACA Advance Cofidential Report ACR No.L4R31 "Profile Drag Coefficients of Conventional and Low Drag Airfoils as Obtained in Flight, dated May 1944.

The Drag build up most often used for the analysis of the P-51B/D were wind tunnel tests performed at CALCIT, namely GALCIT Report 390, dated 2-1-44, "Wind Tunnel Tests on a 1/4 Scale Model of the North American Model NA-102 Airplane Equipped with Running Propellers

3. You state, "the Spit wing despite 3% lower T/C barely edges out the Mustang wing at Vmach>.6 in CDtotal Vs Vmach. Less than 1%." I didn't understand this, could you rephrase? What is Vmach? Is it the critical mach number?

No. The Plots cited are for CDm vs Mach No. For NAA Analysis, the plotted V is expressed in form of Mach no (i.e. 0.5M - which varies with TAS and temperature (T=f(altitude). The value for CDm are required as a multiplier of Parasite Drag components to account for differences between low speed wind tunnel value for Cdo at SL throughout the range of speeds and altitudes for Performance Analysis.

The radiator

So I want to distinguish between two values: gross cooling drag (assuming the radiator installation produced no thrust) and net cooling drag (net of the thrust produced by the Meredith effect).

1. First, just with respect to gross cooling drag (so let's ignore the Meredith effect), how did the P-51's radiator installation compare to the Spitfire Mk IX? Ie, when you say "zero cooling drag" do you mean gross or net?

I haven't seen the Drag build up for Spitfire Parasite Drag. The only gross cooling drag value I have seen are from the NACA testing of P-51B-1 in full scale wind tunnel at Langley. I presented those by reference in my Bastard Stepchild book, as 'Underslung Duct Inlet Uncovered" = 0.0011 @RN=6.19x10^6.

Source Advanced Confidential Report ACR L5A30NACA "Summary of Drag Results from Recent Langley Full Scale Tunnel Tests of Army and Navy Airplanes" dated Feb 1945

2. Second, I've seen conflicting claims on whether the the Meredith effect accounted for any significant thrust. What's the best source on whether it in fact did produce thrust and if so, how much?

Lednicer presented a very strong approach to settling that question via VSAERO CFD model in which he conlcuded that while thrust was achieved, it was not enough to 100% cancel out the internal duct pressure and friction drag. IIRC he concluded that based on a modelling of a 51D at 0.5M at 15000 feet, the associated final drag penalty was 29 pounds for the radiator/oil cooler pressure drag combined with internal friction drag- less Meridith Thrust at that airspeed.

Lednicer also presented a Drag vs Temp plot showing that it was linear as f(exhaust scoop exit Temp T). His assumption was 170F at exhaust for the above bounday conditions.

However, IMO the example was somewhat poorly approached for two reasons. First, his choice of fast cruise = 0.5 M at 15000 feet translates to 360mph - which is not as strenuous on the radiator/oil cooling temperatures as even 61" MP at 3000 RPM which drives an 8800 pound P-51B-1 to about 425mph at 15,000 feet. So, Lednicer's model is tooling along about 60mph below top speed. The second factor was an unknown choice for exit temp. Exit temp varies with exit scoop opening from flush position. At 15K, the 1650-3 and -7 are not delivering peak HP in the blower shift range.

I would invite Lednicer (or Gilchrist) to re-cast the model in conformance with P-51B-1 Flight tests of May 1943. P-51B Performance Test

Below Paragraph 2. of detailed report there is a note that when the exit coolant shutter was fully opened (from +1.5 in from flush) the airspeed dropped from 349IAS to 325 IAS (about 425 mph TAS to 395 mph TAS). That is a drop of 30mph ---------> which when comparing THP1=BHPx550/(425x1.467); THP2=BHPx550/(395x 1.467).
where '550' is Hp conversion factor for Ft.-Lbs/sec,, and 1.467 ft/mph is conversion necessary to render THP to thrust in lbs. BHP at 61in MP at 18000 feet =~1500BHp .
Solve as you wish. Resultant increase in drag = ~100pounds just due to opening exit shutter wide open.

The Fuselage

1. Can you expand on the use of conical sections or provide a link to learn more? Why do they result in low drag?

Thanks so much!

I would draw your attention to the Text/teaching materials developed by Weebe but not likely to be found on internet. I have a very rare original "NA57-548 Master Lines Manual, North American Aviation, Inc. by Robert Kurt Weebe, who was NAA Loftsman, Engineering Group Leader March 1940-1966.

Simply stated, the science of devloping 3-D objects using second degree curve/conics to deveop aerodynamically sound airframes exhibiting constant increasing, but minimum velocity gradients along the freestream flow.


By contrast the very high velocity gradient of flow impinging on round, blunt object to a smaller area region downstream results in a decreased velocity gradient and often seperation of boundary layer. Think radial engine cowl, P-40/Tempest nose structure, or airframe with bumps and scoops...

 

[thunderbird]

I believe all of difference in performance was due the difference paid to excrescence drag on the P-51, or the drag caused by details. At the beginning of the war, excrescence drag was around fifty percent of the base drag. Airplane performances were far worse than what was predicted by wind tunnel tests. Drag due to gaps, to mismatches, to forward facing steps, drag due to skin waviness, drag due to protruding shapes, fasteners, drag due to skin beat up by installing solid metal rivets, flush or otherwise, etcetera. Much of which could be avoided by paying attention to detail design, by tooling that eliminated built in bumps, by better assembly practices that didn't beat up the exterior surfaces.

Now the Spitfire was an exceptionally clean design for its time, with a base drag coefficient that was nearly half the base drag coefficient of the P-38 Lightning (or the Bf109 for that matter). And much better than any of the early American fighters.

But as you might have discovered, the P-51's base drag coefficient was even better.

My theory is that NA paid attention to the details in the P-51 manufacture. After all, there was no chance to achieve any improvement in laminar flow if the surfaces were lumpy and bumpy. And the P-51 development occured after NACA started studying sources of excrescence drag on American fighters.

The laminar wing, the Meredith effect cooling ducts also contributed to the P-51's improved performance.

Interesting enough, the laminar flow wings didn't help the P-51's dive speed as its thicker wing meant the P-51 started seeing drag rise due to compressibility earlier than the Spitfires thinner wings.

 

[SaparotRob]

I believe all of difference in performance was due the difference paid to excrescence drag on the P-51, or the drag caused by details. At the beginning of the war, excrescence drag was around fifty percent of the base drag. Airplane performances were far worse than what was predicted by wind tunnel tests. Drag due to gaps, to mismatches, to forward facing steps, drag due to skin waviness, drag due to protruding shapes, fasteners, drag due to skin beat up by installing solid metal rivets, flush or otherwise, etcetera. Much of which could be avoided by paying attention to detail design, by tooling that eliminated built in bumps, by better assembly practices that didn't beat up the exterior surfaces.

Now the Spitfire was an exceptionally clean design for its time, with a base drag coefficient that was nearly half the base drag coefficient of the P-38 Lightning (or the Bf109 for that matter). And much better than any of the early American fighters.

But as you might have discovered, the P-51's base drag coefficient was even better.

My theory is that NA paid attention to the details in the P-51 manufacture. After all, there was no chance to achieve any improvement in laminar flow if the surfaces were lumpy and bumpy. And the P-51 development occured after NACA started studying sources of excrescence drag on American fighters.

The laminar wing, the Meredith effect cooling ducts also contributed to the P-51's improved performance.

Interesting enough, the laminar flow wings didn't help the P-51's dive speed as its thicker wing meant the P-51 started seeing drag rise due to compressibility earlier than the Spitfires thinner wings.

Great post!

 

[emu27]

drag due to skin waviness

Saw a Spitfire and Mustang parked side by side and the light was such that the skin waviness was most evident on the Spitfire, Mustang on the other hand was as smooth as.