P-51D, Tempest, LA-7 Airfoil Comparisons

[Tilt]

Flap splits seem  pretty much in line with the joint.

btw

The front view on this sketch shows the engine exhaust vanes open either side of the fuselage. 2-3 o'clock and 9-10 o'clock.

When these were open on the La5FN (together with the front cooling vanes open to allow full engine cooling air) max speed was compromised by upto 45 to 50 km/hr and best turning time increased by upto 1.5/2 secs.

Despite the La7's better total aerodynamics (drag wise) a similar effect must of occurred.

In effect WEP was not just increasing revs it was also feathering the entire engine cooling system for minimum drag.

Lavochkin pilots truely had to fly with one eye always on the cylinder head temperature guage.

[Stoney]

Finished the Reynolds Numbers tonight.  After this, I'll start compiling the lift/drag polars.  There will be quite a few.

Reynolds Numbers were generated using the formula:

Rn=p*V*l/u

Where:

p = Density in slugs
V = Velocity in feet/sec
l  = Characteristic length (used Mean Aerodynamic Chord)
u = viscosity in slugs

Method is to compare both the P-51D and LA-7 airfoils at 5 different speeds, at 5 different altitudes.  

Speeds:

150 MPH
200 MPH
250 MPH
300 MPH
350 MPH
400 MPH

Altitudes:

Sea Level
5,000 feet
10,000 feet
15,000 feet
20,000 feet

I couldn't find the MAC for the Tempest, so I didn't include it.  Given the number of polars I'm going to have to generate, I probably did myself a favor ...  Once I have all of the polars generated, I'll compare coefficients of lift, and drag, then compute lift/drag and plot that as well.  These will not be exact, as I don't know how to accurately scale these airfoils to what they would be at the MAC for each aircraft, based on airfoil taper for both.  So, what will be represented is the 23015 (15% thickness) for the LA-7 (which is probably close to the % thickness at the MAC), and the BL 17.5 (17.5% thickness) for the P-51D.  The Pony airfoil will probably be 2-3% thicker than it is in real life, so I expect slightly higher lift and drag coefficients.  Since I'm not espousing anything with this test, merely creating some data for comparison, the accuracy of the Pony airfoil should be close enough for us.  The polars will be generated using XFoil96, using the Reynolds Numbers shown on the chart below, and their corresponding Mach numbers for the altitudes shown.  Altitudes are assumed to be standard day.  Roughness in XFoil will be set to 9.

The chart posted below shows each speed block, broken down by altitude, showing density, velocity, length, and viscosity for each condition.  Two columns display the computed Reynolds numbers, at the far right, one each for the P-51D and LA-7.



Once the polars are generated, I'll be back.

[Stoney]

1st Results:







For some reason, I was getting some faulty numbers for the P-51 above 15 degrees angle of attack.  So, I left off everything above that number for both aircraft.  Obviously, its easy to see the considerable advantage the P-51 airfoil possesses at cruise alphas.  To reduce the amount of work for the comparison (this took me 3 hours), I'm going to pick an altitude for each speed.  So, the 200 mph test will be at 5K, the 250 mph test at 10K, the 300 mph test at 15K, and the 350 mph test at 20K.  I'll leave off the 400 mph test unless there's either some trend that justifies it, or on request.

[Tilt]

 La7 Cy, Cx at various AoA  at sea level at 40.5m/sec

[Tilt]

Any one know what this is for???



Pitching moment v AoA ?

[moot]

I could ask for a translation if no one else does it first..

[Tilt]

Where is fariz when you want him?

[Stoney]

Any one know what this is for???


Pitching moment v AoA ?

That's what it looks like to me.  Its presented differently than I'm used to seeing it, say like the typical NACA graphs, but I'd say pitching moment.

[DPQ5]

they need to make the p51 turn as it did in WW2. But if they did everybody would be flying them and i would get shot down more then i do allready 

[gripen]

Stoney,
You have calculated profiledrag polars while Tilt's data shows the polar of the entire plane so these datasets are not comparable.

[Tilt]

Stoney,
You have calculated profiledrag polars while Tilt's data shows the polar of the entire plane so these datasets are not comparable.

Except to show how far the two can be apart?

[gripen]

Tilt,
The profile drag coefficient is a two dimensional (2D) drag coefficient of a wing section while the the polar you posted shows three dimensional (3D) drag coefficient of the entire airframe. To give some picture how far these are from each other, you can read from your chart that at Cl 0 (Cy) the La-7 had Cd value (Cx) around 0,025 while the theoretical 2D drag coefficient of the NACA 23015 wing section (at MAC) is around 0,006. There are formulas to convert 2D data to 3D using wing geometry but even that does not make the data directly comparable because the effects of the fuselage and tail surfaces are still missing.

[Stoney]

Stoney,
You have calculated profiledrag polars while Tilt's data shows the polar of the entire plane so these datasets are not comparable.

At what point did I represent that I was doing anything other than "airfoil comparisons"?  Tilt added some of those charts to flesh some things out, and that's ok. 

[gripen]

Stoney,
I'm not claiming that you have stated otherwise, sorry if you got such impression. However, these are more or less theoretical values; in practice in flight measured profile drag coefficients of the P-51 were far higher than wind tunnel measured. See the following NACA report on the XP-51:

http://hdl.handle.net/2060/19790073957

To give some idea about the difference in profile drag coefficients at Cl 0,15:

Your data: about 0,0045

From the report:

Unfinished: 0,0070
Factory Finish: 0,0063
Factory Finish + sanding the insigna: 0,0059
Special smoothing by filling and sanding: 0,0053

So even in the best case with special finish, the profile drag coefficient was nearly 20% higher than theoretical wind tunnel value in the case of the P-51. Note that in the case of the NACA 23000 series profiles there probably was very similar difference; in the case of the Fw 190A-5 the Germans measured profile drag coefficient 0,0089 at Cl 0,2 while theoretical value was 0,0067 (in the case of the Bf 109B they measured 0,0101 vs 0,0068 at Cl 0,2).

Notable thing is that in the case of profile of the Mustang, so called "laminar flow bucket" disapeared with standard finish ie the polar shape was in practice very similar with older profiles.

In other words my opinion is that your comparison is more or less theoretical and valid only in the perfect conditions. Tilt's approach by using polars of the entire airframe is much better for comparing airframes.

[Tilt]

Tilt's approach by using polars of the entire airframe is much better for comparing airframes.

It was just some data I had . I have  insufficient knowledge to form an "approach". But I see what you mean    I think Stoney was going thru the theoretical approach to form a comparison rather than a model of reality.  But then I really do not know what I am talking about here.................