NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63-210 AIRFOIL (n63210-il) Reynolds number: 100,000 Max Cl/Cd: 46.66 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63210-il-100000.txt Download as CSV file: xf-n63210-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-210 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4652 0.09208 0.08751 -0.0194 1.0000 0.1467
-9.250 -0.4470 0.08783 0.08324 -0.0176 1.0000 0.1524
-9.000 -0.5520 0.09036 0.08557 -0.0198 1.0000 0.1458
-8.750 -0.5424 0.08670 0.08193 -0.0188 1.0000 0.1531
-8.500 -0.5625 0.08227 0.07761 -0.0239 1.0000 0.1602
-8.250 -0.5491 0.07902 0.07437 -0.0218 1.0000 0.1684
-8.000 -0.5866 0.07352 0.06896 -0.0300 1.0000 0.1750
-7.750 -0.5673 0.07069 0.06618 -0.0267 1.0000 0.1855
-7.500 -0.5701 0.06687 0.06238 -0.0277 1.0000 0.1972
-7.250 -0.5686 0.06382 0.05935 -0.0275 1.0000 0.2132
-6.750 -0.5938 0.04111 0.03423 -0.0386 1.0000 0.0922
-6.500 -0.5765 0.03760 0.03049 -0.0374 1.0000 0.0888
-6.250 -0.5601 0.03376 0.02632 -0.0361 1.0000 0.0837
-6.000 -0.5403 0.03175 0.02338 -0.0339 1.0000 0.0775
-5.750 -0.5213 0.02941 0.02074 -0.0326 1.0000 0.0772
-5.500 -0.5029 0.02681 0.01794 -0.0314 1.0000 0.0786
-5.250 -0.4830 0.02512 0.01604 -0.0299 1.0000 0.0787
-5.000 -0.4629 0.02364 0.01439 -0.0284 1.0000 0.0786
-4.750 -0.4430 0.02221 0.01287 -0.0270 1.0000 0.0792
-4.500 -0.4236 0.02099 0.01164 -0.0256 1.0000 0.0804
-4.250 -0.4045 0.01999 0.01067 -0.0242 1.0000 0.0824
-4.000 -0.3857 0.01914 0.00984 -0.0228 1.0000 0.0853
-3.750 -0.3668 0.01844 0.00911 -0.0215 1.0000 0.0893
-3.500 -0.3477 0.01764 0.00838 -0.0205 1.0000 0.0975
-3.250 -0.3274 0.01695 0.00772 -0.0198 1.0000 0.1106
-3.000 -0.3088 0.01383 0.00690 -0.0197 1.0000 0.5082
-2.750 -0.2963 0.01387 0.00735 -0.0160 1.0000 0.6539
-2.500 -0.2821 0.01414 0.00767 -0.0132 1.0000 0.7036
-2.250 -0.2664 0.01466 0.00830 -0.0099 0.9964 0.7606
-2.000 -0.2472 0.01550 0.00926 -0.0058 0.9857 0.8201
-1.750 -0.2293 0.01606 0.00984 -0.0017 0.9750 0.8626
-1.500 -0.2007 0.01632 0.01002 -0.0010 0.9658 0.8894
-1.250 -0.1613 0.01645 0.00999 -0.0033 0.9576 0.9067
-1.000 -0.1228 0.01649 0.00993 -0.0056 0.9490 0.9229
-0.750 -0.0702 0.01659 0.00989 -0.0107 0.9430 0.9366
-0.500 -0.0212 0.01662 0.00984 -0.0153 0.9352 0.9498
-0.250 0.0408 0.01666 0.00979 -0.0224 0.9302 0.9598
0.000 0.1041 0.01665 0.00972 -0.0298 0.9250 0.9687
0.250 0.1702 0.01654 0.00959 -0.0378 0.9193 0.9756
0.500 0.2354 0.01635 0.00940 -0.0455 0.9146 0.9831
0.750 0.2945 0.01613 0.00922 -0.0523 0.9067 0.9921
1.000 0.3560 0.01579 0.00895 -0.0593 0.9005 1.0000
1.250 0.3799 0.01574 0.00893 -0.0594 0.8877 1.0000
1.500 0.4011 0.01575 0.00897 -0.0589 0.8749 1.0000
1.750 0.4203 0.01583 0.00908 -0.0578 0.8621 1.0000
2.000 0.4380 0.01596 0.00928 -0.0563 0.8495 1.0000
2.250 0.4553 0.01614 0.00949 -0.0547 0.8370 1.0000
2.500 0.4735 0.01631 0.00970 -0.0530 0.8246 1.0000
2.750 0.4944 0.01642 0.00986 -0.0514 0.8121 1.0000
3.000 0.5170 0.01648 0.01001 -0.0500 0.7991 1.0000
3.250 0.5384 0.01648 0.01008 -0.0482 0.7820 1.0000
3.500 0.5612 0.01586 0.00946 -0.0451 0.7572 1.0000
3.750 0.5825 0.01502 0.00857 -0.0413 0.7235 1.0000
4.000 0.6026 0.01425 0.00772 -0.0376 0.6785 1.0000
4.250 0.6235 0.01388 0.00725 -0.0348 0.6274 1.0000
4.500 0.6448 0.01382 0.00711 -0.0327 0.5662 1.0000
4.750 0.6472 0.01560 0.00717 -0.0278 0.2419 1.0000
5.000 0.6553 0.01846 0.00888 -0.0255 0.1020 1.0000
5.250 0.6747 0.01972 0.01012 -0.0242 0.0900 1.0000
5.500 0.6955 0.02089 0.01125 -0.0231 0.0834 1.0000
5.750 0.7174 0.02232 0.01258 -0.0222 0.0786 1.0000
6.000 0.7424 0.02350 0.01381 -0.0216 0.0739 1.0000
6.250 0.7689 0.02499 0.01530 -0.0212 0.0716 1.0000
6.500 0.7966 0.02674 0.01710 -0.0210 0.0700 1.0000
6.750 0.8248 0.02877 0.01923 -0.0208 0.0692 1.0000
7.000 0.8523 0.03100 0.02166 -0.0204 0.0691 1.0000
7.250 0.8784 0.03336 0.02438 -0.0198 0.0699 1.0000
7.500 0.9011 0.03617 0.02773 -0.0186 0.0712 1.0000
7.750 0.9219 0.03891 0.03081 -0.0178 0.0703 1.0000
8.000 0.9370 0.04287 0.03549 -0.0161 0.0732 1.0000
8.250 0.9513 0.04767 0.04072 -0.0148 0.0777 1.0000
8.500 0.9611 0.05480 0.04849 -0.0132 0.0925 1.0000
10.500 0.8305 0.11135 0.10695 -0.0275 0.1632 1.0000
10.750 0.7691 0.11917 0.11455 -0.0403 0.1562 1.0000
11.000 0.6480 0.12249 0.11821 -0.0356 0.1605 1.0000
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Polar data table (+)
Polar graphs
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