9V-SPJ
Topic Author
Posts: 667
Joined: Thu Dec 21, 2000 1:51 pm

Naca Airfoil Data

Sat Aug 27, 2005 5:41 am

Hi all,
Where can I find a graph showing angle of attack vs. lift coefficient for a NACA 0016 airfoil?
Thanks

9V-SPJ
 
air2gxs
Posts: 1443
Joined: Mon Jun 18, 2001 1:29 pm

RE: Naca Airfoil Data

Sat Aug 27, 2005 6:06 am

This what're looking for?

www.eaa1000.av.org/technicl/onedesaf/1desaf.htm

I just typed "naca 0016" into yahoo.

[Edited 2005-08-26 23:07:33]
 
9V-SPJ
Topic Author
Posts: 667
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RE: Naca Airfoil Data

Sat Aug 27, 2005 7:58 am

Thanks for the link, but unfortunately it doesn't have a angle of attack vs. lift coefficient for the graph.
Thanks anyway

9V-SPJ
 
phollingsworth
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RE: Naca Airfoil Data

Sat Aug 27, 2005 9:56 pm

Quoting 9V-SPJ (Thread starter):
Hi all,
Where can I find a graph showing angle of attack vs. lift coefficient for a NACA 0016 airfoil?
Thanks

One source is "Theory of Wing Sections" by Abbott and VonDoenhoff

What you want will be in Appendix IV. It is a Dover book and is only about $15.00.

I don't know of any online sources off-hand
 
flybyguy
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RE: Naca Airfoil Data

Sun Aug 28, 2005 11:21 am

For Christ-sake man! You are at GeorgiaTech! They have an excellent aero engineering program and no doubt an excellent library system.

Any aerodynamics reference book should have a slurry of plots of interest.

And I doubt a school like GeorgiaTech is lacking in those. Just ask your librarian or your professor to find which reference book best suits your needs.
"Are you a pretender... or a thoroughbred?!" - Professor Matt Miller
 
User avatar
zeke
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RE: Naca Airfoil Data

Sun Aug 28, 2005 12:08 pm

NACA 0016 data, sorry dont know how to format the text

Re = 100000
á Cl Cd Cm 0.25 TU TL SU SL L/D
[°] [-] [-] [-] [-] [-] [-] [-] [-]
0.0 0.000 0.01460 -0.000 0.477 0.477 0.981 0.981 0.000
1.0 0.123 0.01466 -0.002 0.414 0.541 0.978 0.981 8.393
2.0 0.246 0.01506 -0.004 0.353 0.606 0.976 0.981 16.310
3.0 0.367 0.01569 -0.005 0.292 0.666 0.969 0.980 23.420
4.0 0.488 0.01661 -0.007 0.233 0.718 0.962 0.979 29.377
5.0 0.606 0.01795 -0.009 0.169 0.763 0.950 0.977 33.756
6.0 0.721 0.01960 -0.011 0.115 0.800 0.931 0.974 36.782
7.0 0.830 0.02175 -0.012 0.068 0.830 0.901 0.972 38.146
8.0 0.930 0.02420 -0.014 0.050 0.851 0.853 0.972 38.438
9.0 1.018 0.02790 -0.015 0.037 0.873 0.776 0.969 36.505
10.0 1.086 0.03456 -0.016 0.030 0.887 0.650 0.971 31.436
11.0 1.130 0.04760 -0.017 0.023 0.902 0.471 0.971 23.743
12.0 1.141 0.07292 -0.016 0.020 0.912 0.228 0.973 15.641
13.0 1.168 0.09395 -0.015 0.017 0.921 0.111 0.974 12.436
14.0 1.203 0.11112 -0.014 0.016 0.931 0.061 0.975 10.829
15.0 1.237 0.12590 -0.014 0.015 0.937 0.043 0.977 9.823
16.0 1.263 0.14178 -0.014 0.014 0.942 0.036 0.978 8.909
17.0 1.280 0.15946 -0.015 0.012 0.948 0.033 0.978 8.028
18.0 1.286 0.17705 -0.015 0.011 0.953 0.027 0.979 7.264
19.0 1.285 0.19758 -0.016 0.010 0.958 0.028 0.979 6.502
20.0 1.274 0.21889 -0.016 0.009 0.961 0.025 0.981 5.820

Re = 200000
á Cl Cd Cm 0.25 TU TL SU SL L/D
[°] [-] [-] [-] [-] [-] [-] [-] [-]
0.0 0.000 0.01140 -0.000 0.477 0.477 0.988 0.988 0.000
1.0 0.123 0.01154 -0.002 0.414 0.541 0.986 0.989 10.685
2.0 0.246 0.01182 -0.004 0.353 0.606 0.983 0.989 20.813
3.0 0.368 0.01234 -0.005 0.292 0.666 0.980 0.989 29.842
4.0 0.489 0.01316 -0.007 0.233 0.718 0.976 0.989 37.185
5.0 0.608 0.01435 -0.009 0.169 0.763 0.966 0.987 42.371
6.0 0.724 0.01585 -0.011 0.115 0.800 0.954 0.985 45.698
7.0 0.835 0.01770 -0.013 0.068 0.830 0.933 0.982 47.204
8.0 0.940 0.01956 -0.014 0.050 0.851 0.904 0.982 48.068
9.0 1.036 0.02205 -0.016 0.037 0.873 0.859 0.979 46.990
10.0 1.120 0.02548 -0.017 0.030 0.887 0.795 0.980 43.975
11.0 1.188 0.03112 -0.018 0.023 0.902 0.699 0.979 38.176
12.0 1.235 0.04031 -0.019 0.020 0.912 0.578 0.979 30.633
13.0 1.262 0.05454 -0.020 0.017 0.921 0.441 0.979 23.142
14.0 1.270 0.07626 -0.019 0.016 0.931 0.287 0.979 16.657
15.0 1.276 0.09979 -0.018 0.015 0.937 0.171 0.980 12.787
16.0 1.285 0.12078 -0.018 0.014 0.942 0.107 0.981 10.641
17.0 1.296 0.13804 -0.018 0.012 0.948 0.082 0.981 9.386
18.0 1.297 0.15634 -0.018 0.011 0.953 0.061 0.981 8.295
19.0 1.295 0.17415 -0.018 0.010 0.958 0.059 0.981 7.434
20.0 1.282 0.19319 -0.019 0.009 0.961 0.053 0.983 6.638

Re = 300000
á Cl Cd Cm 0.25 TU TL SU SL L/D
[°] [-] [-] [-] [-] [-] [-] [-] [-]
0.0 0.000 0.00996 -0.000 0.477 0.477 0.990 0.990 0.000
1.0 0.123 0.01005 -0.002 0.414 0.541 0.990 0.990 12.270
2.0 0.246 0.01029 -0.004 0.353 0.606 0.989 0.990 23.941
3.0 0.369 0.01075 -0.005 0.292 0.666 0.986 0.990 34.300
4.0 0.490 0.01149 -0.007 0.233 0.718 0.982 0.990 42.643
5.0 0.609 0.01250 -0.009 0.169 0.763 0.977 0.990 48.762
6.0 0.726 0.01386 -0.011 0.115 0.800 0.967 0.990 52.406
7.0 0.838 0.01577 -0.013 0.068 0.830 0.949 0.989 53.150
8.0 0.944 0.01752 -0.014 0.050 0.851 0.923 0.987 53.881
9.0 1.042 0.01967 -0.016 0.037 0.873 0.887 0.985 52.969
10.0 1.130 0.02238 -0.017 0.030 0.887 0.837 0.985 50.500
11.0 1.204 0.02662 -0.019 0.023 0.902 0.763 0.983 45.228
12.0 1.259 0.03301 -0.020 0.020 0.912 0.668 0.983 38.147
13.0 1.296 0.04288 -0.021 0.017 0.921 0.558 0.983 30.215
14.0 1.315 0.05741 -0.021 0.016 0.931 0.438 0.982 22.908
15.0 1.323 0.07590 -0.021 0.015 0.937 0.325 0.982 17.433
16.0 1.322 0.09798 -0.021 0.014 0.942 0.226 0.983 13.496
17.0 1.321 0.11915 -0.020 0.012 0.948 0.162 0.983 11.087
18.0 1.315 0.14015 -0.020 0.011 0.953 0.118 0.983 9.383
19.0 1.308 0.15817 -0.020 0.010 0.958 0.100 0.983 8.268
20.0 1.293 0.17772 -0.021 0.009 0.961 0.086 0.984 7.276

Re = 400000
á Cl Cd Cm 0.25 TU TL SU SL L/D
[°] [-] [-] [-] [-] [-] [-] [-] [-]
0.0 0.000 0.00902 -0.000 0.477 0.477 0.991 0.991 0.000
1.0 0.123 0.00909 -0.002 0.414 0.541 0.990 0.991 13.564
2.0 0.246 0.00907 -0.004 0.353 0.606 0.991 0.991 27.161
3.0 0.369 0.00946 -0.006 0.292 0.666 0.990 0.991 38.993
4.0 0.490 0.01042 -0.007 0.233 0.718 0.987 0.991 47.062
5.0 0.610 0.01135 -0.009 0.169 0.763 0.982 0.991 53.759
6.0 0.728 0.01268 -0.011 0.115 0.800 0.977 0.991 57.374
7.0 0.840 0.01434 -0.013 0.068 0.830 0.962 0.990 58.579
8.0 0.947 0.01608 -0.014 0.050 0.851 0.940 0.990 58.874
9.0 1.046 0.01816 -0.016 0.037 0.873 0.906 0.989 57.585
10.0 1.135 0.02061 -0.018 0.030 0.887 0.859 0.989 55.084
11.0 1.213 0.02418 -0.019 0.023 0.902 0.796 0.986 50.148
12.0 1.272 0.02940 -0.020 0.020 0.912 0.715 0.986 43.262
13.0 1.312 0.03739 -0.021 0.017 0.921 0.617 0.986 35.095
14.0 1.337 0.04899 -0.022 0.016 0.931 0.511 0.984 27.285
15.0 1.349 0.06417 -0.022 0.015 0.937 0.408 0.985 21.017
16.0 1.349 0.08339 -0.022 0.014 0.942 0.310 0.985 16.172
17.0 1.344 0.10390 -0.022 0.012 0.948 0.234 0.985 12.933
18.0 1.333 0.12584 -0.022 0.011 0.953 0.173 0.985 10.590
19.0 1.321 0.14529 -0.022 0.010 0.958 0.142 0.985 9.092
20.0 1.303 0.16544 -0.022 0.009 0.961 0.117 0.986 7.876

Re = 500000
á Cl Cd Cm 0.25 TU TL SU SL L/D
[°] [-] [-] [-] [-] [-] [-] [-] [-]
0.0 0.000 0.00835 -0.000 0.477 0.477 0.991 0.991 0.000
1.0 0.123 0.00843 -0.002 0.414 0.541 0.991 0.991 14.643
2.0 0.246 0.00840 -0.004 0.353 0.606 0.991 0.991 29.343
3.0 0.369 0.00878 -0.006 0.292 0.666 0.991 0.991 42.019
4.0 0.491 0.00935 -0.007 0.233 0.718 0.990 0.991 52.468
5.0 0.611 0.01057 -0.009 0.169 0.763 0.987 0.991 57.784
6.0 0.728 0.01182 -0.011 0.115 0.800 0.981 0.991 61.620
7.0 0.842 0.01342 -0.013 0.068 0.830 0.971 0.991 62.746
8.0 0.950 0.01506 -0.014 0.050 0.851 0.953 0.990 63.058
9.0 1.049 0.01692 -0.016 0.037 0.873 0.922 0.990 62.033
10.0 1.139 0.01926 -0.018 0.030 0.887 0.877 0.990 59.167
11.0 1.218 0.02259 -0.019 0.023 0.902 0.817 0.989 53.906
12.0 1.280 0.02711 -0.020 0.020 0.912 0.745 0.989 47.214
13.0 1.323 0.03402 -0.021 0.017 0.921 0.655 0.988 38.889
14.0 1.350 0.04400 -0.022 0.016 0.931 0.557 0.986 30.686
15.0 1.364 0.05721 -0.023 0.015 0.937 0.460 0.987 23.849
16.0 1.366 0.07433 -0.023 0.014 0.942 0.365 0.987 18.375
17.0 1.360 0.09384 -0.023 0.012 0.948 0.285 0.987 14.492
18.0 1.347 0.11478 -0.023 0.011 0.953 0.218 0.986 11.737
19.0 1.332 0.13536 -0.023 0.010 0.958 0.175 0.986 9.839
20.0 1.311 0.15597 -0.023 0.009 0.961 0.144 0.987 8.408
We are addicted to our thoughts. We cannot change anything if we cannot change our thinking – Santosh Kalwar
 
L-188
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RE: Naca Airfoil Data

Sun Aug 28, 2005 1:42 pm

I seem to recall seeing some PDF copies of 1910's on NCAC reports on a NASA website somewhere. But for the life of me I can't remember where.

But there where some reports of Air-tunnel testing of airfoils on it.
OBAMA-WORST PRESIDENT EVER....Even SKOORB would be better.
 
vikkyvik
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RE: Naca Airfoil Data

Sun Aug 28, 2005 5:31 pm

I'm sorry, but I agree with Flybyguy on this one....If I had a question like this, my first source would be the AE department at school (and I'm an AE major as well). Actually, my first source would be the first aerospace engineering book I ever purchased, which was by Anderson (I think it was Fundamentals of Flight, or something like that). It had a bunch of CL graphs in the back, and while I don't know if the 0016 was among them, it very well might have been.

No offense intended, but you have to use the resources that a university affords you. A.net, while good for some questions, is not the one-stop for all college folk.

~Vik
I'm watching Jeopardy. The category is worst Madonna songs. "This one from 1987 is terrible".
 
liedetectors
Posts: 323
Joined: Thu Jul 21, 2005 10:44 am

RE: Naca Airfoil Data

Mon Aug 29, 2005 2:15 am

Back at the University of Illinois, we used a program called xfoil to get airfoil data. Its a bit cumbersome to use since u have to load an airfoil data file, then for each run (you can vary the AOA), you will need to record the data into excel or some other spread sheet file. THen you can make your own Cl-alpha curves. It is a UNIX based program so watch out. Very handy tool once you figure it out. Ask your lowspeed aerodynamics prof, he should know about it. Mark Drella from MIT made the program if i do recall. He and our prof at UIUC are some of the top aerodynamisists in the world.
If it was said by us, then it must be true.
 
aeroweanie
Posts: 1577
Joined: Fri Dec 03, 2004 9:33 pm

RE: Naca Airfoil Data

Mon Aug 29, 2005 4:53 am

Quoting Liedetectors (Reply 8):
Its a bit cumbersome to use since u have to load an airfoil data file, then for each run (you can vary the AOA), you will need to record the data into excel or some other spread sheet file.

Xfoil can create all of the NACA 4-digit and 5-digit airfoils internally.

Quoting Liedetectors (Reply 8):
It is a UNIX based program so watch out.

No, there is a Windows version available for free on the MIT website.

Quoting Liedetectors (Reply 8):
Mark Drella from MIT made the program if i do recall.

Yes, Mark wrote it, but his last name is Drela, not Drella. He is a prof at MIT.
 
liedetectors
Posts: 323
Joined: Thu Jul 21, 2005 10:44 am

RE: Naca Airfoil Data

Mon Aug 29, 2005 7:59 am

Quoting AeroWeanie (Reply 9):
No, there is a Windows version available for free on the MIT website.

Is there really? can you send me a link? My bust on the spelling.
If it was said by us, then it must be true.
 
phollingsworth
Posts: 635
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RE: Naca Airfoil Data

Mon Aug 29, 2005 10:05 am

Quoting Flybyguy (Reply 4):
For Christ-sake man! You are at GeorgiaTech! They have an excellent aero engineering program and no doubt an excellent library system.

Any aerodynamics reference book should have a slurry of plots of interest.

And I doubt a school like GeorgiaTech is lacking in those. Just ask your librarian or your professor to find which reference book best suits your needs.

I didn't even notice that I would have told him to stop by my office I have the data there, of course it wouldn't do a whole lot of good right now as I am out of town on travel for a few days.
 
aeroweanie
Posts: 1577
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RE: Naca Airfoil Data

Mon Aug 29, 2005 10:37 am

Quoting Liedetectors (Reply 10):
Is there really? can you send me a link? My bust on the spelling.

Here you go: http://raphael.mit.edu/xfoil/

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