CLARK Y AIRFOIL (clarky-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: CLARK Y AIRFOIL (clarky-il) Reynolds number: 50,000 Max Cl/Cd: 29.62 at α=9.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-clarky-il-50000.txt Download as CSV file: xf-clarky-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: CLARK Y AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3410 0.10621 0.09925 -0.0263 1.0000 0.2711
-8.250 -0.3291 0.10240 0.09548 -0.0251 1.0000 0.2802
-8.000 -0.3587 0.10280 0.09609 -0.0238 1.0000 0.2885
-7.750 -0.3342 0.09804 0.09130 -0.0223 1.0000 0.3003
-7.500 -0.3487 0.09645 0.08986 -0.0203 1.0000 0.3079
-7.250 -0.3539 0.09474 0.08824 -0.0178 1.0000 0.3191
-7.000 -0.3507 0.09182 0.08538 -0.0156 1.0000 0.3276
-6.750 -0.3882 0.09242 0.08620 -0.0108 1.0000 0.3366
-6.500 -0.3717 0.08866 0.08244 -0.0090 1.0000 0.3477
-6.250 -0.4070 0.08840 0.08239 -0.0044 1.0000 0.3557
-6.000 -0.4206 0.08715 0.08124 -0.0011 1.0000 0.3691
-5.750 -0.4114 0.08412 0.07824 0.0019 1.0000 0.3819
-5.500 -0.4177 0.08215 0.07635 0.0052 1.0000 0.3970
-5.250 -0.4281 0.08049 0.07478 0.0087 1.0000 0.4150
-5.000 -0.4512 0.07946 0.07387 0.0119 1.0000 0.4355
-4.750 -0.4331 0.07637 0.07079 0.0164 1.0000 0.4583
-4.500 -0.4314 0.05863 0.05189 -0.0290 1.0000 0.2245
-4.250 -0.4061 0.05169 0.04446 -0.0335 1.0000 0.1898
-4.000 -0.3795 0.04648 0.03857 -0.0369 1.0000 0.1744
-3.750 -0.3552 0.04308 0.03475 -0.0383 1.0000 0.1689
-3.500 -0.3276 0.03982 0.03076 -0.0400 1.0000 0.1648
-3.250 -0.3046 0.03807 0.02876 -0.0402 1.0000 0.1675
-3.000 -0.2805 0.03649 0.02684 -0.0405 1.0000 0.1710
-2.750 -0.2551 0.03494 0.02488 -0.0409 1.0000 0.1738
-2.500 -0.2291 0.03365 0.02309 -0.0411 1.0000 0.1773
-2.250 -0.2061 0.03262 0.02207 -0.0410 1.0000 0.1837
-2.000 -0.1821 0.03190 0.02114 -0.0410 1.0000 0.1953
-1.750 -0.1587 0.03128 0.02045 -0.0408 1.0000 0.2103
-1.500 -0.1351 0.03065 0.01990 -0.0406 1.0000 0.2310
-1.250 -0.1094 0.03013 0.01949 -0.0407 1.0000 0.2753
-1.000 -0.0812 0.02923 0.01921 -0.0413 1.0000 0.3787
-0.750 -0.0716 0.02623 0.01894 -0.0359 1.0000 1.0000
-0.500 -0.0468 0.02692 0.01880 -0.0368 1.0000 1.0000
-0.250 -0.0042 0.02806 0.01942 -0.0410 0.9917 1.0000
0.000 0.0506 0.02938 0.02029 -0.0473 0.9751 1.0000
0.250 0.1051 0.03060 0.02116 -0.0532 0.9579 1.0000
0.500 0.1570 0.03168 0.02194 -0.0585 0.9411 1.0000
0.750 0.1995 0.03254 0.02259 -0.0620 0.9246 1.0000
1.000 0.2387 0.03332 0.02321 -0.0648 0.9084 1.0000
1.250 0.2766 0.03407 0.02382 -0.0673 0.8923 1.0000
1.500 0.3130 0.03477 0.02441 -0.0693 0.8759 1.0000
1.750 0.3483 0.03541 0.02497 -0.0711 0.8590 1.0000
2.000 0.3833 0.03603 0.02551 -0.0726 0.8421 1.0000
2.250 0.4186 0.03658 0.02602 -0.0740 0.8249 1.0000
2.500 0.4544 0.03706 0.02645 -0.0753 0.8077 1.0000
2.750 0.4906 0.03746 0.02683 -0.0766 0.7907 1.0000
3.000 0.5275 0.03775 0.02713 -0.0777 0.7739 1.0000
3.250 0.5643 0.03796 0.02734 -0.0785 0.7574 1.0000
3.500 0.6003 0.03811 0.02751 -0.0792 0.7413 1.0000
3.750 0.6355 0.03821 0.02763 -0.0796 0.7255 1.0000
4.000 0.6685 0.03834 0.02780 -0.0797 0.7100 1.0000
4.250 0.6981 0.03860 0.02810 -0.0793 0.6946 1.0000
4.500 0.7256 0.03897 0.02850 -0.0787 0.6795 1.0000
4.750 0.7506 0.03952 0.02909 -0.0779 0.6647 1.0000
5.000 0.7734 0.04025 0.02989 -0.0770 0.6504 1.0000
5.250 0.7952 0.04107 0.03076 -0.0761 0.6366 1.0000
5.500 0.8193 0.04177 0.03152 -0.0753 0.6240 1.0000
5.750 0.8676 0.04061 0.03046 -0.0759 0.6151 1.0000
6.000 0.8800 0.04212 0.03204 -0.0743 0.6010 1.0000
6.250 0.8907 0.04383 0.03382 -0.0727 0.5874 1.0000
6.500 0.9033 0.04541 0.03548 -0.0713 0.5743 1.0000
6.750 0.9267 0.04608 0.03626 -0.0702 0.5625 1.0000
7.000 0.9915 0.04315 0.03346 -0.0710 0.5528 1.0000
7.250 1.0017 0.04470 0.03512 -0.0690 0.5386 1.0000
7.500 1.0164 0.04584 0.03639 -0.0672 0.5244 1.0000
7.750 1.0478 0.04535 0.03601 -0.0660 0.5094 1.0000
8.000 1.1029 0.04267 0.03340 -0.0657 0.4921 1.0000
8.250 1.1433 0.04141 0.03218 -0.0650 0.4741 1.0000
8.500 1.1591 0.04217 0.03306 -0.0628 0.4566 1.0000
8.750 1.1777 0.04275 0.03375 -0.0609 0.4383 1.0000
9.000 1.2049 0.04271 0.03379 -0.0594 0.4185 1.0000
9.250 1.2442 0.04201 0.03301 -0.0589 0.3967 1.0000
9.500 1.2528 0.04349 0.03464 -0.0562 0.3771 1.0000
9.750 1.2746 0.04420 0.03538 -0.0545 0.3559 1.0000
10.000 1.3042 0.04488 0.03594 -0.0538 0.3353 1.0000
10.250 1.3144 0.04680 0.03799 -0.0515 0.3190 1.0000
10.500 1.3280 0.04857 0.03985 -0.0496 0.3038 1.0000
10.750 1.3409 0.05037 0.04176 -0.0477 0.2902 1.0000
11.000 1.3177 0.05428 0.04606 -0.0431 0.2829 1.0000
11.250 1.3370 0.05612 0.04800 -0.0420 0.2723 1.0000
11.500 1.2972 0.06109 0.05330 -0.0370 0.2691 1.0000
11.750 1.2377 0.06803 0.06039 -0.0328 0.2687 1.0000
12.000 1.1120 0.08579 0.07807 -0.0355 0.2728 1.0000
12.250 1.3618 0.06228 0.05459 -0.0327 0.2297 1.0000
12.500 1.3213 0.06787 0.06046 -0.0287 0.2300 1.0000
12.750 1.3277 0.06776 0.06037 -0.0262 0.2160 1.0000
13.000 1.3716 0.06211 0.05435 -0.0240 0.1874 1.0000
13.250 1.3285 0.06839 0.06104 -0.0213 0.1888 1.0000
13.500 1.2920 0.07486 0.06775 -0.0204 0.1885 1.0000
13.750 1.2584 0.08181 0.07483 -0.0208 0.1874 1.0000
14.500 1.2881 0.08254 0.07524 -0.0147 0.1321 1.0000
14.750 1.2556 0.08980 0.08275 -0.0162 0.1318 1.0000
15.000 0.8641 0.18509 0.17714 -0.0828 0.3321 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CLARK Y AIRFOIL (clarky-il)