NACA 66(1)-212 (naca661212-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 66(1)-212 (naca661212-il) Reynolds number: 200,000 Max Cl/Cd: 46.79 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca661212-il-200000-n5.txt Download as CSV file: xf-naca661212-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.5184 0.09064 0.08698 -0.0510 1.0000 0.0188
-11.000 -0.5888 0.06776 0.06399 -0.0663 1.0000 0.0171
-10.750 -0.6174 0.06187 0.05796 -0.0675 1.0000 0.0169
-10.500 -0.6394 0.05851 0.05451 -0.0661 1.0000 0.0170
-10.250 -0.6595 0.05542 0.05131 -0.0642 0.9932 0.0170
-10.000 -0.6595 0.05208 0.04779 -0.0657 0.9803 0.0173
-9.750 -0.6478 0.05026 0.04586 -0.0669 0.9718 0.0177
-9.500 -0.6355 0.04816 0.04360 -0.0680 0.9646 0.0185
-9.250 -0.6326 0.04424 0.03934 -0.0675 0.9555 0.0188
-9.000 -0.6265 0.04041 0.03508 -0.0668 0.9479 0.0191
-8.750 -0.6189 0.03716 0.03143 -0.0653 0.9398 0.0194
-8.500 -0.6073 0.03421 0.02807 -0.0639 0.9334 0.0199
-8.250 -0.5945 0.03175 0.02521 -0.0623 0.9267 0.0204
-8.000 -0.5767 0.02981 0.02282 -0.0611 0.9220 0.0216
-7.750 -0.5590 0.02844 0.02106 -0.0597 0.9163 0.0223
-7.500 -0.5393 0.02612 0.01848 -0.0589 0.9117 0.0228
-7.250 -0.5164 0.02451 0.01669 -0.0585 0.9080 0.0234
-7.000 -0.4931 0.02332 0.01536 -0.0581 0.9042 0.0240
-6.750 -0.4700 0.02230 0.01422 -0.0576 0.8996 0.0248
-6.500 -0.4460 0.02149 0.01329 -0.0572 0.8957 0.0261
-6.250 -0.4211 0.02068 0.01235 -0.0568 0.8927 0.0276
-6.000 -0.3963 0.01981 0.01137 -0.0565 0.8896 0.0286
-5.750 -0.3725 0.01904 0.01052 -0.0559 0.8857 0.0294
-5.500 -0.3488 0.01840 0.00979 -0.0553 0.8822 0.0301
-5.250 -0.3283 0.01733 0.00870 -0.0544 0.8791 0.0317
-5.000 -0.3066 0.01670 0.00805 -0.0535 0.8764 0.0335
-4.750 -0.2855 0.01627 0.00761 -0.0526 0.8722 0.0362
-4.500 -0.2637 0.01585 0.00713 -0.0518 0.8685 0.0385
-4.250 -0.2410 0.01546 0.00664 -0.0511 0.8655 0.0404
-4.000 -0.2189 0.01497 0.00612 -0.0502 0.8630 0.0444
-3.750 -0.1955 0.01465 0.00576 -0.0497 0.8605 0.0499
-3.500 -0.1731 0.01435 0.00547 -0.0490 0.8570 0.0576
-3.250 -0.1513 0.01393 0.00517 -0.0482 0.8539 0.0853
-3.000 -0.1392 0.01262 0.00474 -0.0462 0.8505 0.2907
-2.750 -0.1343 0.01106 0.00458 -0.0424 0.8475 0.5958
-2.500 -0.1138 0.01122 0.00509 -0.0402 0.8449 0.7211
-2.250 -0.0916 0.01157 0.00548 -0.0386 0.8418 0.7618
-2.000 -0.0677 0.01191 0.00582 -0.0372 0.8392 0.7833
-1.750 -0.0427 0.01214 0.00602 -0.0362 0.8370 0.7965
-1.500 -0.0174 0.01233 0.00617 -0.0354 0.8350 0.8080
-1.250 0.0088 0.01246 0.00625 -0.0348 0.8331 0.8166
-1.000 0.0361 0.01246 0.00619 -0.0348 0.8313 0.8212
-0.750 0.0607 0.01256 0.00629 -0.0344 0.8282 0.8243
-0.500 0.0865 0.01261 0.00632 -0.0342 0.8256 0.8273
-0.250 0.1131 0.01262 0.00630 -0.0343 0.8232 0.8301
0.000 0.1403 0.01260 0.00625 -0.0345 0.8209 0.8332
0.250 0.1682 0.01258 0.00622 -0.0347 0.8190 0.8357
0.500 0.1964 0.01259 0.00622 -0.0350 0.8175 0.8373
0.750 0.2215 0.01266 0.00632 -0.0347 0.8141 0.8394
1.000 0.2464 0.01271 0.00641 -0.0344 0.8100 0.8420
1.250 0.2736 0.01270 0.00642 -0.0344 0.8066 0.8448
1.500 0.3023 0.01265 0.00638 -0.0347 0.8039 0.8476
1.750 0.3302 0.01262 0.00638 -0.0350 0.8008 0.8501
2.000 0.3537 0.01270 0.00655 -0.0343 0.7955 0.8522
2.250 0.3813 0.01264 0.00655 -0.0342 0.7910 0.8542
2.500 0.4084 0.01246 0.00641 -0.0339 0.7830 0.8567
2.750 0.4381 0.01210 0.00608 -0.0337 0.7722 0.8593
3.000 0.4637 0.01183 0.00584 -0.0329 0.7572 0.8625
3.250 0.4866 0.01160 0.00565 -0.0316 0.7354 0.8655
3.500 0.5088 0.01128 0.00530 -0.0299 0.6942 0.8679
3.750 0.5264 0.01125 0.00476 -0.0270 0.5625 0.8709
4.000 0.5266 0.01237 0.00503 -0.0219 0.4032 0.8760
4.250 0.5283 0.01357 0.00555 -0.0177 0.2597 0.8811
4.500 0.5315 0.01480 0.00613 -0.0138 0.1225 0.8861
4.750 0.5438 0.01564 0.00663 -0.0115 0.0620 0.8914
5.000 0.5620 0.01610 0.00708 -0.0100 0.0504 0.8958
5.250 0.5799 0.01652 0.00754 -0.0084 0.0445 0.9001
5.500 0.5983 0.01695 0.00802 -0.0069 0.0401 0.9053
5.750 0.6135 0.01752 0.00863 -0.0050 0.0368 0.9109
6.000 0.6290 0.01798 0.00917 -0.0029 0.0345 0.9168
6.500 0.6621 0.01903 0.01038 0.0005 0.0303 0.9309
6.750 0.6787 0.01972 0.01111 0.0020 0.0287 0.9391
7.000 0.6962 0.02066 0.01209 0.0033 0.0274 0.9471
7.250 0.7186 0.02183 0.01331 0.0037 0.0263 0.9544
7.500 0.7460 0.02273 0.01433 0.0032 0.0254 0.9616
7.750 0.7758 0.02369 0.01540 0.0023 0.0241 0.9688
8.000 0.8066 0.02471 0.01652 0.0010 0.0226 0.9774
8.250 0.8385 0.02596 0.01790 -0.0005 0.0217 0.9900
8.500 0.8652 0.02723 0.01926 -0.0010 0.0210 1.0000
8.750 0.8907 0.02868 0.02082 -0.0014 0.0204 1.0000
9.000 0.9164 0.03064 0.02290 -0.0018 0.0199 1.0000
9.250 0.9393 0.03313 0.02563 -0.0018 0.0195 1.0000
9.500 0.9562 0.03503 0.02786 -0.0007 0.0193 1.0000
9.750 0.9687 0.03695 0.03015 0.0010 0.0187 1.0000
10.000 0.9783 0.03905 0.03259 0.0030 0.0180 1.0000
10.250 0.9833 0.04147 0.03533 0.0054 0.0176 1.0000
10.500 0.9834 0.04407 0.03824 0.0084 0.0173 1.0000
10.750 0.9789 0.04689 0.04136 0.0115 0.0171 1.0000
11.000 0.9709 0.04981 0.04456 0.0146 0.0169 1.0000
11.250 0.9597 0.05301 0.04802 0.0174 0.0168 1.0000
11.500 0.9453 0.05646 0.05172 0.0199 0.0167 1.0000
11.750 0.9276 0.06037 0.05587 0.0218 0.0168 1.0000
12.000 0.9075 0.06473 0.06044 0.0229 0.0167 1.0000
12.250 0.8843 0.06980 0.06572 0.0230 0.0168 1.0000
12.500 0.8597 0.07557 0.07169 0.0217 0.0169 1.0000
12.750 0.8312 0.08285 0.07914 0.0186 0.0170 1.0000
13.000 0.7995 0.09259 0.08903 0.0127 0.0172 1.0000
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Polar data table (+)
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