NASA SC(2)-0610 AIRFOIL (sc20610-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA SC(2)-0610 AIRFOIL (sc20610-il) Reynolds number: 200,000 Max Cl/Cd: 43.12 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20610-il-200000-n5.txt Download as CSV file: xf-sc20610-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0610 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-12.750 -0.7273 0.08578 0.08161 -0.0357 1.0000 0.0228
-12.500 -0.8075 0.06322 0.05872 -0.0529 1.0000 0.0221
-12.250 -0.8398 0.05583 0.05107 -0.0569 1.0000 0.0220
-12.000 -0.8652 0.05065 0.04566 -0.0580 1.0000 0.0220
-11.750 -0.8875 0.04692 0.04172 -0.0567 1.0000 0.0220
-11.500 -0.9040 0.04348 0.03803 -0.0555 1.0000 0.0222
-11.250 -0.9101 0.03996 0.03417 -0.0552 1.0000 0.0224
-11.000 -0.9084 0.03687 0.03070 -0.0545 1.0000 0.0227
-10.750 -0.9006 0.03417 0.02763 -0.0539 1.0000 0.0230
-10.500 -0.8881 0.03190 0.02501 -0.0533 1.0000 0.0234
-10.250 -0.8724 0.03018 0.02316 -0.0528 1.0000 0.0238
-10.000 -0.8545 0.02889 0.02176 -0.0524 1.0000 0.0243
-9.750 -0.8354 0.02771 0.02049 -0.0520 1.0000 0.0248
-9.500 -0.8154 0.02655 0.01921 -0.0516 1.0000 0.0255
-9.250 -0.7946 0.02547 0.01798 -0.0512 1.0000 0.0264
-9.000 -0.7731 0.02441 0.01674 -0.0509 1.0000 0.0276
-8.750 -0.7510 0.02338 0.01552 -0.0505 1.0000 0.0286
-8.500 -0.7287 0.02242 0.01456 -0.0503 1.0000 0.0297
-8.250 -0.7054 0.02164 0.01375 -0.0502 1.0000 0.0308
-8.000 -0.6817 0.02086 0.01289 -0.0500 1.0000 0.0322
-7.750 -0.6573 0.02009 0.01202 -0.0499 1.0000 0.0339
-7.500 -0.6323 0.01930 0.01119 -0.0501 1.0000 0.0356
-7.250 -0.6069 0.01877 0.01067 -0.0503 1.0000 0.0379
-7.000 -0.5811 0.01827 0.01010 -0.0504 1.0000 0.0409
-6.750 -0.5544 0.01766 0.00948 -0.0509 1.0000 0.0439
-6.500 -0.5276 0.01721 0.00902 -0.0512 1.0000 0.0473
-6.250 -0.5005 0.01680 0.00855 -0.0516 1.0000 0.0512
-6.000 -0.4723 0.01631 0.00810 -0.0523 1.0000 0.0559
-5.750 -0.4443 0.01600 0.00773 -0.0528 1.0000 0.0610
-5.500 -0.4145 0.01553 0.00732 -0.0539 1.0000 0.0672
-5.250 -0.3818 0.01518 0.00694 -0.0554 0.9991 0.0746
-5.000 -0.3476 0.01479 0.00660 -0.0572 0.9980 0.0842
-4.750 -0.3131 0.01443 0.00629 -0.0591 0.9969 0.0977
-4.500 -0.2777 0.01404 0.00599 -0.0613 0.9958 0.1178
-4.000 -0.2029 0.01290 0.00546 -0.0670 0.9947 0.2471
-3.750 -0.1612 0.01192 0.00521 -0.0713 0.9953 0.4087
-3.500 -0.1260 0.01158 0.00558 -0.0731 0.9947 0.5619
-3.250 -0.0937 0.01169 0.00574 -0.0739 0.9924 0.6089
-3.000 -0.0621 0.01181 0.00587 -0.0746 0.9897 0.6324
-2.750 -0.0306 0.01198 0.00607 -0.0752 0.9871 0.6524
-2.500 0.0024 0.01221 0.00628 -0.0762 0.9849 0.6738
-2.250 0.0324 0.01247 0.00658 -0.0763 0.9818 0.6875
-2.000 0.0597 0.01272 0.00687 -0.0758 0.9778 0.6975
-1.750 0.0916 0.01289 0.00703 -0.0765 0.9751 0.7074
-1.500 0.1228 0.01306 0.00724 -0.0769 0.9726 0.7129
-1.250 0.1560 0.01311 0.00730 -0.0781 0.9694 0.7188
-1.000 0.1895 0.01308 0.00728 -0.0792 0.9638 0.7235
-0.750 0.2295 0.01297 0.00722 -0.0814 0.9589 0.7260
-0.500 0.2679 0.01273 0.00701 -0.0832 0.9490 0.7284
-0.250 0.3127 0.01232 0.00664 -0.0861 0.9381 0.7308
0.000 0.3527 0.01190 0.00624 -0.0879 0.9205 0.7333
0.250 0.3926 0.01150 0.00583 -0.0897 0.8938 0.7361
0.500 0.4295 0.01126 0.00556 -0.0910 0.8621 0.7392
0.750 0.4620 0.01115 0.00541 -0.0913 0.8154 0.7408
1.000 0.4904 0.01139 0.00515 -0.0903 0.6851 0.7426
1.250 0.5058 0.01262 0.00536 -0.0874 0.4802 0.7446
1.500 0.5256 0.01366 0.00571 -0.0861 0.3304 0.7470
1.750 0.5490 0.01446 0.00603 -0.0855 0.2259 0.7497
2.000 0.5749 0.01509 0.00631 -0.0855 0.1565 0.7525
2.250 0.6026 0.01559 0.00659 -0.0858 0.1194 0.7553
2.500 0.6282 0.01600 0.00692 -0.0853 0.0997 0.7570
2.750 0.6540 0.01638 0.00728 -0.0849 0.0865 0.7589
3.000 0.6801 0.01676 0.00766 -0.0845 0.0771 0.7610
3.250 0.7061 0.01722 0.00807 -0.0842 0.0695 0.7634
3.500 0.7331 0.01759 0.00849 -0.0840 0.0632 0.7661
3.750 0.7597 0.01812 0.00897 -0.0840 0.0578 0.7688
4.000 0.7877 0.01851 0.00940 -0.0841 0.0528 0.7715
4.250 0.8122 0.01907 0.00994 -0.0835 0.0485 0.7730
4.500 0.8377 0.01950 0.01047 -0.0830 0.0449 0.7748
4.750 0.8629 0.02001 0.01102 -0.0825 0.0419 0.7768
5.000 0.8874 0.02069 0.01170 -0.0819 0.0393 0.7792
5.250 0.9130 0.02127 0.01238 -0.0815 0.0365 0.7819
5.500 0.9389 0.02182 0.01295 -0.0813 0.0342 0.7848
5.750 0.9640 0.02264 0.01379 -0.0810 0.0325 0.7874
6.000 0.9878 0.02345 0.01474 -0.0801 0.0310 0.7892
6.250 1.0116 0.02429 0.01570 -0.0794 0.0296 0.7912
6.500 1.0354 0.02512 0.01660 -0.0787 0.0285 0.7935
6.750 1.0590 0.02603 0.01756 -0.0781 0.0276 0.7960
7.000 1.0824 0.02723 0.01887 -0.0775 0.0268 0.7987
7.250 1.1065 0.02844 0.02031 -0.0770 0.0257 0.8015
7.500 1.1287 0.02961 0.02169 -0.0761 0.0247 0.8037
7.750 1.1498 0.03073 0.02298 -0.0751 0.0239 0.8058
8.000 1.1703 0.03195 0.02438 -0.0740 0.0233 0.8082
8.250 1.1902 0.03321 0.02578 -0.0729 0.0228 0.8110
8.500 1.2092 0.03466 0.02736 -0.0719 0.0224 0.8141
8.750 1.2259 0.03671 0.02965 -0.0706 0.0221 0.8171
9.000 1.2377 0.03927 0.03268 -0.0684 0.0217 0.8193
9.250 1.2448 0.04235 0.03625 -0.0658 0.0214 0.8216
9.500 1.2457 0.04602 0.04044 -0.0627 0.0211 0.8240
9.750 1.2389 0.05029 0.04521 -0.0591 0.0208 0.8266
10.000 1.2233 0.05500 0.05039 -0.0551 0.0206 0.8294
10.250 1.1967 0.05958 0.05534 -0.0503 0.0205 0.8325
10.500 1.1641 0.06514 0.06125 -0.0471 0.0205 0.8348
10.750 1.1239 0.07264 0.06910 -0.0469 0.0206 0.8370
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Polar data table (+)
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