NACA 65(1)-212 (naca651212-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 65(1)-212 (naca651212-il) Reynolds number: 200,000 Max Cl/Cd: 52.86 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca651212-il-200000-n5.txt Download as CSV file: xf-naca651212-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65(1)-212
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.5657 0.08916 0.08550 -0.0376 1.0000 0.0186
-11.000 -0.5998 0.07419 0.07050 -0.0485 1.0000 0.0180
-10.750 -0.6338 0.06512 0.06129 -0.0542 1.0000 0.0177
-10.500 -0.6680 0.05826 0.05423 -0.0565 1.0000 0.0174
-10.250 -0.6827 0.05506 0.05091 -0.0561 1.0000 0.0176
-10.000 -0.7051 0.05142 0.04710 -0.0539 1.0000 0.0175
-9.750 -0.7115 0.04935 0.04491 -0.0517 1.0000 0.0178
-9.500 -0.7174 0.04680 0.04219 -0.0494 1.0000 0.0182
-9.250 -0.7232 0.04413 0.03931 -0.0466 1.0000 0.0185
-9.000 -0.7305 0.04150 0.03645 -0.0429 1.0000 0.0187
-8.750 -0.7164 0.03748 0.03199 -0.0436 0.9892 0.0190
-8.500 -0.6957 0.03388 0.02791 -0.0448 0.9806 0.0195
-8.250 -0.6714 0.03088 0.02443 -0.0458 0.9732 0.0203
-8.000 -0.6452 0.02880 0.02186 -0.0466 0.9655 0.0214
-7.750 -0.6188 0.02659 0.01932 -0.0473 0.9582 0.0221
-7.500 -0.5906 0.02470 0.01726 -0.0483 0.9517 0.0228
-7.250 -0.5628 0.02336 0.01579 -0.0489 0.9444 0.0234
-7.000 -0.5343 0.02215 0.01442 -0.0495 0.9374 0.0242
-6.750 -0.5075 0.02112 0.01325 -0.0497 0.9296 0.0251
-6.500 -0.4806 0.02028 0.01230 -0.0499 0.9221 0.0266
-6.250 -0.4552 0.01946 0.01136 -0.0496 0.9140 0.0278
-6.000 -0.4300 0.01862 0.01040 -0.0493 0.9064 0.0285
-5.750 -0.4074 0.01771 0.00943 -0.0485 0.8981 0.0294
-5.500 -0.3856 0.01680 0.00849 -0.0478 0.8907 0.0308
-5.250 -0.3635 0.01617 0.00783 -0.0470 0.8825 0.0322
-5.000 -0.3401 0.01566 0.00727 -0.0464 0.8755 0.0347
-4.750 -0.3168 0.01521 0.00673 -0.0458 0.8676 0.0371
-4.500 -0.2934 0.01469 0.00611 -0.0452 0.8611 0.0397
-4.250 -0.2698 0.01421 0.00560 -0.0446 0.8539 0.0437
-4.000 -0.2450 0.01387 0.00516 -0.0442 0.8474 0.0498
-3.750 -0.2209 0.01340 0.00474 -0.0438 0.8409 0.0646
-3.500 -0.1990 0.01262 0.00429 -0.0432 0.8340 0.1395
-3.250 -0.1830 0.01105 0.00373 -0.0423 0.8282 0.3787
-3.000 -0.1633 0.01034 0.00373 -0.0410 0.8214 0.5587
-2.750 -0.1388 0.01026 0.00383 -0.0401 0.8160 0.6293
-2.500 -0.1143 0.01037 0.00403 -0.0390 0.8100 0.6809
-2.250 -0.0889 0.01048 0.00417 -0.0381 0.8042 0.7064
-2.000 -0.0620 0.01052 0.00414 -0.0377 0.7997 0.7192
-1.750 -0.0343 0.01050 0.00405 -0.0378 0.7938 0.7267
-1.500 -0.0067 0.01048 0.00398 -0.0377 0.7883 0.7312
-1.250 0.0212 0.01046 0.00388 -0.0377 0.7838 0.7363
-1.000 0.0491 0.01045 0.00383 -0.0379 0.7780 0.7422
-0.750 0.0767 0.01045 0.00380 -0.0378 0.7730 0.7466
-0.500 0.1046 0.01045 0.00375 -0.0379 0.7687 0.7517
-0.250 0.1325 0.01045 0.00374 -0.0380 0.7629 0.7577
0.000 0.1599 0.01046 0.00375 -0.0380 0.7581 0.7620
0.250 0.1879 0.01047 0.00374 -0.0380 0.7540 0.7673
0.500 0.2157 0.01050 0.00378 -0.0382 0.7483 0.7733
0.750 0.2430 0.01052 0.00383 -0.0380 0.7435 0.7779
1.000 0.2709 0.01055 0.00386 -0.0381 0.7395 0.7836
1.250 0.2984 0.01059 0.00395 -0.0382 0.7340 0.7893
1.500 0.3256 0.01063 0.00404 -0.0380 0.7291 0.7944
1.750 0.3535 0.01067 0.00410 -0.0381 0.7250 0.8006
2.000 0.3806 0.01073 0.00424 -0.0380 0.7197 0.8062
2.250 0.4077 0.01079 0.00437 -0.0379 0.7145 0.8119
2.500 0.4357 0.01084 0.00445 -0.0380 0.7101 0.8186
2.750 0.4617 0.01089 0.00463 -0.0376 0.7028 0.8241
3.000 0.4882 0.01088 0.00466 -0.0372 0.6918 0.8310
3.250 0.5128 0.01081 0.00462 -0.0362 0.6713 0.8372
3.500 0.5363 0.01073 0.00449 -0.0349 0.6376 0.8440
3.750 0.5582 0.01077 0.00440 -0.0335 0.5874 0.8512
4.000 0.5793 0.01096 0.00443 -0.0320 0.5258 0.8588
4.250 0.5920 0.01173 0.00461 -0.0292 0.3933 0.8673
4.500 0.5984 0.01322 0.00530 -0.0261 0.2283 0.8775
4.750 0.6078 0.01448 0.00597 -0.0235 0.1066 0.8872
5.000 0.6236 0.01528 0.00654 -0.0216 0.0601 0.8978
5.250 0.6422 0.01583 0.00705 -0.0202 0.0463 0.9092
5.500 0.6615 0.01630 0.00757 -0.0188 0.0407 0.9215
5.750 0.6815 0.01678 0.00815 -0.0176 0.0369 0.9350
6.000 0.7012 0.01752 0.00892 -0.0166 0.0333 0.9512
6.250 0.7289 0.01809 0.00960 -0.0171 0.0308 0.9683
6.500 0.7554 0.01882 0.01040 -0.0177 0.0288 1.0000
6.750 0.7747 0.01959 0.01120 -0.0169 0.0274 1.0000
7.000 0.7932 0.02047 0.01210 -0.0159 0.0263 1.0000
7.250 0.8103 0.02168 0.01331 -0.0148 0.0253 1.0000
7.500 0.8305 0.02281 0.01448 -0.0140 0.0246 1.0000
7.750 0.8526 0.02377 0.01553 -0.0134 0.0239 1.0000
8.000 0.8750 0.02476 0.01663 -0.0130 0.0229 1.0000
8.250 0.8974 0.02580 0.01776 -0.0125 0.0219 1.0000
8.500 0.9199 0.02698 0.01904 -0.0121 0.0211 1.0000
8.750 0.9425 0.02830 0.02048 -0.0118 0.0206 1.0000
9.000 0.9640 0.02969 0.02200 -0.0114 0.0200 1.0000
9.250 0.9848 0.03122 0.02366 -0.0109 0.0196 1.0000
9.500 1.0046 0.03300 0.02559 -0.0104 0.0192 1.0000
9.750 1.0229 0.03538 0.02815 -0.0098 0.0188 1.0000
10.000 1.0356 0.03768 0.03075 -0.0084 0.0184 1.0000
10.250 1.0443 0.03967 0.03308 -0.0065 0.0180 1.0000
10.500 1.0478 0.04181 0.03557 -0.0040 0.0174 1.0000
10.750 1.0473 0.04447 0.03855 -0.0013 0.0171 1.0000
11.000 1.0429 0.04726 0.04166 0.0014 0.0169 1.0000
11.250 1.0346 0.05030 0.04499 0.0038 0.0168 1.0000
11.500 1.0230 0.05364 0.04860 0.0059 0.0167 1.0000
11.750 1.0080 0.05741 0.05263 0.0074 0.0165 1.0000
12.000 0.9907 0.06160 0.05706 0.0082 0.0165 1.0000
12.250 0.9710 0.06635 0.06205 0.0081 0.0165 1.0000
12.500 0.9496 0.07173 0.06764 0.0069 0.0166 1.0000
12.750 0.9247 0.07820 0.07431 0.0044 0.0166 1.0000
13.000 0.8983 0.08576 0.08203 0.0005 0.0168 1.0000
13.250 0.8680 0.09557 0.09201 -0.0058 0.0169 1.0000
13.500 0.8331 0.10976 0.10633 -0.0154 0.0173 1.0000
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Polar data table (+)
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