Long range, high speed Spitfire fighter: the best approach?

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Bill, maybe it would be the best to look at these fight tests, in hope nothing gets lost in translation?

Anyway, reading from the power chart Denniss kindly posted, at sea level, on 'Start-und-Notleistung' power setting, the engine was giving 1780 PS. The D-9 was tested, according to this, 557 km/h (346 mp/h) at SL ('Start-und-Notleistung Normal' setting, the engine gaps were not sealed). The normal take off weight of the D-9 is quoted at 4270 kg (~9405 lbs) by following table.

I've partly translated part of the table, earlier kindly provided by our member Bada, seems that on high speed the A-8/9 have had some 10% greater Cd than D-9? (open the table separately for hi-res)

 
I will crank over the next 24 hours. Offhand the Drag numbers in that chart look awfully high. For example D-9 wing and tail parasite drag per translation (only=.1647+.0440)= .2117 which is ridiculously high (by 15x) compared to what I would expect for Zero lift parasite drag of the wing and tail
 
We might arrive at interesting numbers, with high-speed Cd divided by wing area. Ie. For the D-9, 0.444/18.9 gives 0.0234 (in the ballpark with Spit XI), and A-8/9: 0,485/18.9 gives 0.0256. Don't quote me on this - the Cw (wiederstand coeffizient - drag coefficient) actually has the measuring unit - square meter. Equivalent flat plate?? 0.444 m^2 is equal to 4.779 sq ft.
Help.

edit - Lednicer gives f values for the A and D as being 5.22 and 4.77, respectively. Fw data gives for A-8 0.485 m^2 = 5.22 sq ft. Nice
edit2: German term is 'Wiederstandflache' (CwF) - ie. 'Drag plate'
 
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D/q = Cd*S so, what we see in the rollup data is the "D/q" calculation.. and then Cd follows as Drag Force divided by Area..

which now works..D in pounds, Q in pounds/sq.ft = equivalent square feet of flat plate drag

So (D/q)/S=Cd: .444/18.9 = .002349

I wasn't familiar with the presentation but shoulda noted the "(delta W)/q *m>>2) and recognized it as

(delta D/q) = Cd*S for each major sub component.. too used to English units rather than metric and too old to convert.

q= .5*rho*(V*V); rho = .002377 in slugs ; V*V in fps = 346*1.467*(346*1.467)= 257,639.5 Ft.sq/sec.sq
q= 306 pounds/sq.ft
CL = (W/S)/q = (9405/197)/306.2 = .1559; CL>>2 = (.1559)>>2 = .0243

CDi= (CL)>>2/(Pi*AR*e); assume 'e' = ~.85 and assume about a 5% reduction in Aspect Ratio AR due to wingtip drag.

AR= (wing span)>>2/Wing Area *(1-.05) = ((34.4>>2)/197 )* .95= 5.8

CDi= .0243/(3.14*5.8*.85)= .00157

But CD = CDo+CDi; solving for Cdo -----> = CD-CDi = .02349 - .00157 = .02192 for Zero Lift Parasite Drag via SL dash speed data.

By comparison, the 51 was around .017, the Spit IX about .023 and the 109G about .026+
=
 
Tomo, The charts come from an old RAE report (got to chase it down again) and they have different scales.
I re-scaled them (and made mach adjustments),bu they are a bit of a mess.

I've been meaning to do a tidied version for ages, might do it this weekend.
 

11.45pm UK time is the wrong time to open a post like this and try to understand it.
 
OK finally got these wind tunnel stuff sorted out (sort of).
They are from RAE Report 2535 and are a combination of test on the Spit Mk1, Mustang Mk, Spiteful and Attacker (plus alternate Spiteful windscreens).

They (as the report shows) only a reasonable fit to real life test, but they key thing is that they are a constant test without distortions caused by radiator efficiency, fit and finish, etc.
Therefore they represent the theoretical results of the body/fuselage/tail.

There is a key point for the Spitfire numbers, they had made a mistake in their readings and corrections.
As the article mentions that for comparison between the Spit and all the others you have to correct the mach numbers:
0.6 becomes 0.607, 0.7 becomes 0.712, 0.75 become 0.768, 0.78 become 0.807, 0.8 becomes 0.838.

They say that they are not reliable above mach 0.75, which explains the odd Spit results at CL 0.4

I have put the charts together and (reasonably) matched them.
Because the CL varies with mach I have done 4 charts: mach 0.4, 0.6, 0.7 and 0.75. I'll show them in separate posts
The green vertical lines correct for the Spit mach numbers. I chose to keep the Spit's ones constant and show the corresponding adjusted value for the Mustang.
Blue horizontal lines are for the Mustang and red for the Spit.
The key is where the green and blue/red lines intersect.

Mach 0.4


As we can see the Mustang is better at mach 0.4 at CL 0. Almost identical at CL 0.2 and CL0.4.
 
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Mach 0.7 (note the correction).


Similar picture Mustang better at 0, same at 0.2 and worse at 0.4
 
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Mach 0.75. Again note the mach correction.


Now things change. Mustang slightly better at CL 0, but now worse at Cl 0.2 and much worse at CL 0.4,
 
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Basic overall conclusions, given the levels of accuracy of the measurements and my scaling and selecting, they were very close in most flight regimes of CL and speed. Only at the extremes do any significant differences show.

Note that on my early post about this I hadn't been so systematic about the different mach speeds and tried to 'eyeball' average values.
Definitely got some wrong then.
 
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OldSkeptic said:
Mach 0.75. Again note the mach correction.
View attachment 234081

Now things change. Mustang slightly better at CL 0, but now worse at Cl 0.2 and much worse at CL 0.4,
Click to expand...

Several items interesting about the plot. Question - do you have the originals? When I look at CL=.1 for the Spit the Total Drag CD is LOWER for CL=.1 from .5M to .7M than at CL=0 which should not be the case. In all cases total Drag should be at minimum when aircraft in Zero Lift AoA... (no Induced Drag and Form Drag at minimum)

Drag = Zero Lift Parasite Drag + Induced Drag + Form Drag + Compressibility Drag.

AT CL=0 all these plots reflect increased compressibility (which becomes meaningful above .35M).. The Spit demonstrates no compressibility drag rise (IMO because of thin wing) until the .5M region. The Spit CD rises non-linearly in the .6M to .7M where it enters Onset Drag region (i.e. region where CD increases ~ .002 over short interval of velocity).

Note: The Spit wing had max T/c at ~25% so the velocity gradient from free .stream Vo is higher than the more gradual increase for the Mustang at ~ 45%. It will enter the Mcr profile Sooner but will ultimately have less drag at M>.85 because of the thinner wing.

The Mustang demonstrates a steady but small increase of CD until .6M to .7M where it 'flattens out' (IMO -again the wing) and doesn't reflect onset Drag Rise until .72-.75M.

In ALL flight velocity range @CL=0 the CD of the Mustang is lower to significantly lower (.018@.7M to Spit .023@.7M). This reflects reality for all reported flight tests for level flight speed runs.

At CL=0 (which could only occur for zero G dive at .7M) the real differences are significant. For CL>0 the aircraft are in positive G increasingly high angles of attack where form drag increases dramatically... hence the 'spikes' for each CL increase..

Was this a full scale wind tunnel test? The RN is curiously low.
 
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It is written by someone with a cursory knowledge of aero terms, has some facts and opinions not so well based on facts.

First - the Straight stall speed for the Mustang will occur at CL max ~1.78 to 1.89 depending on GW (w/Flaps) and the CLmax for the 51 is ~ 1.6 for level flight (depending on GW).

The CLmax for the 109 w/Flaps and LE slats is ~ 2.0 for level flight. What is not clear is what effect (trim drag for example), the extra lift for the top wing causes to require increased rudder deflection to overcome natural yaw due to the LE Slats..

What is not stated (and I believe not well understood by the author of that piece) is that the form drag (i.e. BL separation on Fuse and wings, trim drag of deflected rudder/ailerons, etc) related to high AoA flight conditions is extraordinarily high near CLmax as separation is occurring on both the wings and the fuselage. I don't have the data but I suspect that the reason that both the Spit and the P-51 competed well in turn maneuvers with the 109 was that the combined Parasite and Form Drag of the Spit/51 was lower than the 109 in high G maneuvers. Staying power in a turning (or climbing) fight is all about a.) WL and b.) Excess Thrust over Drag.

According to the author, the only reason one could surmise that the 109 got shot down in a maneuvering fight is incompetency on the part of the 109 pilot.

I suppose zjtins has a reason to post the opinion..but that is not particularly evident from prior track record of his posts so far..Do ya suppose Gaston has returned?
 
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Don't apologise - I love the charts and hope to look at the originals.. They are consistent with other P-51 Drag vs M for different CL in that the drag rise for higher CL is showing up in the .65M range as it should.
 
Bill, you are, how sould I express this... to generous for someone undeserving (ie. to the author of the web page, linked at post 154).
The author of the page lists the 109G as a 1941 machine, capable then to do 416 mph (we need to wait for 1942 for the F4 and G-1/2 to do this). Then he claims that Spit V was able to do 400 mph, and P-51D (of 1943!) capable for 395 mph??? Range with drop tanks - 624 miles for 109G, vs. 950 miles for the P-51D?? Rate of climb - 3300 fpm 109G, 1700 fpm P-51D - every time now I'll run out from question marks...
That is from Comparison chart #2, the chart #1 is nowhere to be seen. However, the chart #2 is actually a table.
The caption under a picture there says:

The P-51 Mustang and the Spitfire are often thought of as the mightiest fighters of WWII. They were not. The Me109 was and here are some of its secrets! The SECRET of the Messerschmitt Me109G!
Click to expand...

What a source.
 
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