NACA 66(1)-212 (naca661212-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 66(1)-212 (naca661212-il) Reynolds number: 500,000 Max Cl/Cd: 71.09 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca661212-il-500000.txt Download as CSV file: xf-naca661212-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66(1)-212
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.4064 0.08722 0.08510 -0.0517 1.0000 0.0246
-11.000 -0.4132 0.08219 0.08010 -0.0534 1.0000 0.0248
-10.750 -0.4238 0.07664 0.07458 -0.0555 0.9992 0.0251
-10.500 -0.4496 0.06426 0.06219 -0.0649 0.9930 0.0248
-10.250 -0.4907 0.05279 0.05054 -0.0732 0.9861 0.0244
-10.000 -0.5094 0.04662 0.04421 -0.0768 0.9783 0.0244
-9.750 -0.5197 0.04179 0.03921 -0.0790 0.9693 0.0245
-9.500 -0.5285 0.03847 0.03574 -0.0785 0.9560 0.0248
-8.000 -0.5849 0.02941 0.02428 -0.0602 0.9139 0.0235
-7.750 -0.5665 0.02832 0.02294 -0.0586 0.9082 0.0230
-7.500 -0.5520 0.02502 0.01930 -0.0569 0.9033 0.0233
-7.250 -0.5312 0.02332 0.01734 -0.0558 0.8996 0.0232
-7.000 -0.5091 0.02112 0.01492 -0.0551 0.8956 0.0233
-6.750 -0.4849 0.01918 0.01283 -0.0548 0.8919 0.0238
-6.500 -0.4601 0.01794 0.01151 -0.0545 0.8886 0.0243
-6.250 -0.4357 0.01722 0.01073 -0.0542 0.8856 0.0255
-6.000 -0.4108 0.01659 0.01005 -0.0539 0.8825 0.0267
-5.750 -0.3857 0.01582 0.00922 -0.0536 0.8790 0.0277
-5.500 -0.3609 0.01507 0.00841 -0.0531 0.8756 0.0284
-5.250 -0.3364 0.01450 0.00776 -0.0526 0.8726 0.0293
-5.000 -0.3113 0.01417 0.00736 -0.0522 0.8700 0.0301
-4.750 -0.2931 0.01293 0.00609 -0.0508 0.8666 0.0323
-4.500 -0.2702 0.01248 0.00565 -0.0502 0.8634 0.0345
-4.250 -0.2467 0.01208 0.00521 -0.0496 0.8605 0.0365
-4.000 -0.2224 0.01175 0.00483 -0.0491 0.8580 0.0385
-3.750 -0.1988 0.01135 0.00437 -0.0485 0.8557 0.0412
-3.500 -0.1747 0.01102 0.00404 -0.0480 0.8533 0.0466
-3.250 -0.1496 0.01078 0.00378 -0.0477 0.8504 0.0525
-3.000 -0.1268 0.01029 0.00349 -0.0471 0.8475 0.0974
-2.750 -0.1211 0.00817 0.00291 -0.0444 0.8442 0.4896
-2.500 -0.1044 0.00747 0.00298 -0.0423 0.8419 0.7001
-2.250 -0.0774 0.00755 0.00307 -0.0420 0.8402 0.7311
-2.000 -0.0513 0.00772 0.00328 -0.0415 0.8380 0.7584
-1.750 -0.0253 0.00793 0.00353 -0.0409 0.8357 0.7783
-1.500 0.0012 0.00812 0.00374 -0.0404 0.8334 0.7917
-1.250 0.0281 0.00825 0.00387 -0.0401 0.8311 0.8017
-1.000 0.0553 0.00839 0.00401 -0.0398 0.8292 0.8095
-0.750 0.0824 0.00852 0.00413 -0.0396 0.8276 0.8173
-0.500 0.1102 0.00862 0.00423 -0.0395 0.8261 0.8221
-0.250 0.1382 0.00869 0.00428 -0.0397 0.8242 0.8264
0.000 0.1659 0.00871 0.00430 -0.0400 0.8218 0.8294
0.250 0.1935 0.00870 0.00431 -0.0402 0.8190 0.8311
0.500 0.2215 0.00867 0.00429 -0.0403 0.8157 0.8330
0.750 0.2502 0.00862 0.00423 -0.0405 0.8127 0.8348
1.000 0.2791 0.00863 0.00422 -0.0409 0.8101 0.8367
1.250 0.3063 0.00862 0.00426 -0.0410 0.8064 0.8390
1.500 0.3342 0.00853 0.00418 -0.0411 0.8009 0.8415
1.750 0.3633 0.00840 0.00400 -0.0412 0.7943 0.8438
2.000 0.3899 0.00823 0.00386 -0.0409 0.7854 0.8455
2.250 0.4168 0.00806 0.00371 -0.0405 0.7750 0.8474
2.500 0.4440 0.00792 0.00355 -0.0402 0.7639 0.8494
2.750 0.4698 0.00776 0.00341 -0.0396 0.7475 0.8517
3.000 0.4953 0.00763 0.00326 -0.0390 0.7256 0.8543
3.250 0.5208 0.00759 0.00321 -0.0385 0.7005 0.8572
3.500 0.5431 0.00764 0.00312 -0.0373 0.6433 0.8599
3.750 0.5418 0.00893 0.00348 -0.0318 0.4390 0.8635
4.000 0.5399 0.01059 0.00414 -0.0269 0.2210 0.8679
4.250 0.5456 0.01199 0.00475 -0.0235 0.0610 0.8723
4.500 0.5654 0.01243 0.00514 -0.0222 0.0460 0.8754
4.750 0.5870 0.01272 0.00548 -0.0212 0.0416 0.8784
5.000 0.6068 0.01315 0.00594 -0.0198 0.0374 0.8822
5.250 0.6246 0.01375 0.00659 -0.0182 0.0346 0.8867
5.500 0.6446 0.01411 0.00701 -0.0169 0.0327 0.8905
5.750 0.6631 0.01453 0.00747 -0.0153 0.0304 0.8946
6.000 0.6805 0.01505 0.00801 -0.0137 0.0284 0.8994
6.250 0.6919 0.01602 0.00905 -0.0110 0.0270 0.9047
6.500 0.7084 0.01671 0.00982 -0.0091 0.0260 0.9096
6.750 0.7282 0.01729 0.01046 -0.0078 0.0252 0.9151
7.000 0.7487 0.01799 0.01123 -0.0067 0.0244 0.9201
7.250 0.7702 0.01874 0.01205 -0.0057 0.0235 0.9257
7.500 0.7933 0.01949 0.01285 -0.0052 0.0226 0.9317
7.750 0.8127 0.02011 0.01352 -0.0039 0.0217 0.9386
8.000 0.8395 0.02127 0.01471 -0.0042 0.0209 0.9445
8.250 0.8757 0.02391 0.01747 -0.0062 0.0203 0.9467
8.500 0.9047 0.02723 0.02106 -0.0068 0.0200 0.9509
8.750 0.9213 0.02846 0.02252 -0.0051 0.0198 0.9611
9.000 0.9415 0.03019 0.02453 -0.0043 0.0197 0.9722
9.250 0.9657 0.03241 0.02706 -0.0046 0.0193 0.9836
9.500 0.9838 0.03468 0.02967 -0.0040 0.0188 1.0000
9.750 0.9908 0.03837 0.03370 -0.0020 0.0191 1.0000
10.000 0.9558 0.05426 0.05041 0.0040 0.0264 1.0000
10.250 0.9502 0.05679 0.05317 0.0071 0.0264 1.0000
10.500 0.9387 0.05844 0.05502 0.0112 0.0262 1.0000
10.750 0.9279 0.05778 0.05466 0.0164 0.0243 1.0000
11.000 0.9103 0.06098 0.05803 0.0193 0.0243 1.0000
11.250 0.8913 0.06440 0.06161 0.0215 0.0238 1.0000
11.500 0.8704 0.06839 0.06574 0.0226 0.0236 1.0000
11.750 0.8468 0.07313 0.07061 0.0227 0.0236 1.0000
12.000 0.8210 0.07882 0.07642 0.0213 0.0238 1.0000
12.250 0.7925 0.08603 0.08375 0.0178 0.0238 1.0000
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Polar data table (+)
Polar graphs
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