Information on Boeing 747 Wing

[rbogie]

I posted a thread in a different area with two questions. I got a couple of replies to the first part, but no replies to the second part so I'll start a new thread around the second question here. I am trying to find out data on the Boeing 747 wing, specifically the airfoil and the L/D curves for it in order to approximately calculate the Lift and drag of the wing (for a 10th grade science fair.) I found on a website that the B-747 has an airfoil of "BAC 463 thru 468". What does that mean? Does the wing have varying airfoils from root to tip? What does a BAC 463, 464, 465, 466, 467, 468 look like? What do the L/D curves look like? Other, simpler airplanes use NACA 4 digit or 5 digit airfoils and I can look up the graphs of their LD curves, but I have no idea how to look up Boeing L vs AoA or D vs AoA curves.

 

[rb1957]

i think this is boeing proprietary data, and probably impossible to get a hold of.

maybe you could turn the project around, and reverse engineer the aerodynamic properties based on a flight simulator ?

bare in mind the comments on the previous thread about compressibility effects.

possibly you could get more information on an out of production airplane (say B707, DC8, DC10).

good luck

 

[rbogie]

Well the B-707 uses BAC 310/313/312 airfoil, the B-727 uses BAC 424/425/426/427 airfoil.

Not sure how the FAA certifies airplanes, but I would have thought that they'd file drawings with the Gov't somewhere, that something as basic as the airfoil shape wouldn't be kept secret, do they usually put patents on an airfoil shape? I thought airfoils were generated by a formula.

 

[rb1957]

the FAA probably doesn't have copies of the drawings. they probably have copies of the analysis reports boeing prepared for certification BUT these are covered by non-disclosure agreements. these reports are company confidential and can't be released by the FAA without permission.

i don't think the profile is patented, tho' there might be some design feature in the profile that is unique. you're right rpofiles are generated by a formulae, and that's the secret ! NACA (the precursor to NASA) has a bunch of profiles and L/D and CL data that are available to the public, but the trouble is relating these to the profiles Boeing used.

 

[rbogie]

That's exactly what I was getting at. If a BAC airfoil was the same as a NACA 64a-315 (just an example) then I could look up the L and D graphs for that wing.
Even if not that, then if there was some way of getting the Boeing L and D graphs for analysis would be sufficient.

 

[wes616]

Boeing had released some of the geometry information for the 747-200 for a CFD study. You should be able to use the data from this to get some numbers. There are numerous papers online. Try to look at.

http://aero-comlab.stanford.edu/Papers/jameson.aiaa.03-3438.pdf#search="boeing 747-200 airfoil"

And yes, airfoils change from root to tip.

Wes C.
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No trees were killed in the sending of this message, but a large number of electrons were terribly inconvenienced.

 

[rbogie]

Thanks Wes. I think I'll tell my teenager to switch to an airplane that has a single airfoil that we can look up and find the Lift and Drag curves for.

 

[rb1957]

i think you'll find that most wings vary from tip to root ... sometimes the profile changes, most often the incidence changes (twist, wash-out). but this isn't that hard to account for, you were planning on three sections to define the wing, use an average to these three.

eg ... for stall performance, say your three sections are the same profile but with twist, maybe the tip has lower incidence than the root, maybe the incidence of the tip is equal to the plane's incidence, at the mid-span it might be +1 degree, at the root +2 degree. Thus when the root CL is maximum (stall) the CL at the other two sections can be determined. each section will have a portion of the wing area, this is how to average the three sections together,; so now you'll have the combined wing CL*S.
then Vstall = sqrt((1.1*Wt)*2/(rho*CL*S))

and this is flaps up

 

[rbogie]

I don't understand how you get that 1.1 factor? Is that an actual number or just an example of a factor to account for averaging the CL's? I could see how you would use three CL's in the formula in this way: Sqrt((Wt)*(2/(rho*((CL1*S1)+(CL2*S2)+(CL3*S3))))--I hope I got the correct number of parenthesis.

It may be difficult to actually get the correct surface area for each "Zone" of respective airfoil/CL, but I could probably get an approximation by finding the distance out from the Centerline for each zone,(remember, the reference surface area usually used is the total "Wetted" area, including the extension into the fuselage profile--not just from the wing root outwards), finding what proportion that distance is to the total span and dividing the total area by that fraction. Find the respective Y distance for each zone to get a fraction of total span and multiply that fraction of total Area to get each respective area.

 

[rb1957]

an allowance for horizontal tail (down) lift ... like we talked about in your first thread.

as for your (CLi*Si) ... yeah, I'd associate a nominal area for each section, maybe 20% for the tip, 50% for the mid section and 30% for the root. maybe play with the %ages to see how much difference there is.

 

[Sparweb]

No guarantee that it's right, but they seem so thorough...

http://www.ae.uiuc.edu/m-selig/ads/aircraft.html

Steven Fahey, CET

 

[rbogie]

Yeah, I looked there, that's how I figured out that the B-747 wing uses 6 different airfoils.