DF 101 AIRFOIL (df101-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: DF 101 AIRFOIL (df101-il) Reynolds number: 100,000 Max Cl/Cd: 50.52 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-df101-il-100000.txt Download as CSV file: xf-df101-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: DF 101 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4155 0.10444 0.09946 -0.0281 1.0000 0.1247
-9.500 -0.4123 0.10118 0.09624 -0.0287 1.0000 0.1288
-9.250 -0.4578 0.09874 0.09403 -0.0370 1.0000 0.1340
-9.000 -0.4307 0.09367 0.08891 -0.0332 1.0000 0.1359
-8.750 -0.4205 0.09036 0.08561 -0.0322 1.0000 0.1382
-8.500 -0.4194 0.08721 0.08251 -0.0323 1.0000 0.1413
-7.750 -0.5724 0.05481 0.04990 -0.0477 1.0000 0.0651
-7.500 -0.5807 0.05082 0.04577 -0.0454 1.0000 0.0626
-7.250 -0.6077 0.04434 0.03816 -0.0415 1.0000 0.0558
-7.000 -0.6015 0.04109 0.03468 -0.0392 1.0000 0.0547
-6.750 -0.5958 0.03765 0.03091 -0.0369 1.0000 0.0540
-6.500 -0.5869 0.03465 0.02750 -0.0347 1.0000 0.0537
-6.250 -0.5749 0.03212 0.02452 -0.0325 1.0000 0.0541
-6.000 -0.5609 0.03002 0.02193 -0.0305 1.0000 0.0559
-5.750 -0.5447 0.02776 0.01962 -0.0291 1.0000 0.0592
-5.500 -0.5261 0.02607 0.01771 -0.0276 1.0000 0.0615
-5.250 -0.5061 0.02458 0.01596 -0.0260 1.0000 0.0648
-5.000 -0.4868 0.02310 0.01441 -0.0247 1.0000 0.0709
-4.750 -0.4673 0.02216 0.01336 -0.0232 1.0000 0.0789
-4.500 -0.4488 0.02103 0.01232 -0.0219 1.0000 0.0897
-4.250 -0.4301 0.01999 0.01135 -0.0204 1.0000 0.1078
-4.000 -0.4121 0.01882 0.01059 -0.0190 1.0000 0.1624
-3.750 -0.3947 0.01849 0.01051 -0.0177 1.0000 0.2497
-3.500 -0.3772 0.01853 0.01058 -0.0165 1.0000 0.2870
-3.250 -0.3416 0.01881 0.01087 -0.0187 0.9942 0.3308
-3.000 -0.3027 0.01895 0.01107 -0.0215 0.9858 0.3739
-2.750 -0.2644 0.01888 0.01104 -0.0239 0.9777 0.4087
-2.500 -0.2220 0.01880 0.01100 -0.0271 0.9703 0.4452
-2.250 -0.1879 0.01855 0.01093 -0.0286 0.9609 0.4827
-2.000 -0.1499 0.01834 0.01092 -0.0307 0.9528 0.5355
-1.750 -0.1139 0.01797 0.01095 -0.0322 0.9443 0.6147
-1.500 -0.0812 0.01746 0.01114 -0.0321 0.9364 0.7684
-1.250 0.0535 0.01752 0.01123 -0.0519 0.9426 1.0000
-1.000 0.0891 0.01761 0.01110 -0.0541 0.9311 1.0000
-0.750 0.1295 0.01771 0.01102 -0.0569 0.9211 1.0000
-0.500 0.1802 0.01773 0.01088 -0.0614 0.9136 1.0000
-0.250 0.2162 0.01777 0.01081 -0.0632 0.9020 1.0000
0.000 0.2624 0.01770 0.01065 -0.0667 0.8931 1.0000
0.250 0.3092 0.01749 0.01035 -0.0700 0.8843 1.0000
0.500 0.3428 0.01737 0.01018 -0.0707 0.8717 1.0000
0.750 0.3770 0.01719 0.00995 -0.0714 0.8595 1.0000
1.000 0.4114 0.01693 0.00965 -0.0719 0.8473 1.0000
1.250 0.4452 0.01661 0.00929 -0.0722 0.8354 1.0000
1.500 0.4781 0.01626 0.00891 -0.0721 0.8236 1.0000
1.750 0.5036 0.01610 0.00872 -0.0709 0.8095 1.0000
2.000 0.5287 0.01595 0.00854 -0.0696 0.7951 1.0000
2.250 0.5539 0.01580 0.00838 -0.0683 0.7805 1.0000
2.500 0.5789 0.01567 0.00821 -0.0669 0.7654 1.0000
2.750 0.6031 0.01558 0.00809 -0.0655 0.7494 1.0000
3.000 0.6256 0.01556 0.00806 -0.0638 0.7316 1.0000
3.250 0.6492 0.01552 0.00801 -0.0623 0.7134 1.0000
3.500 0.6736 0.01547 0.00792 -0.0609 0.6948 1.0000
3.750 0.6980 0.01545 0.00784 -0.0595 0.6752 1.0000
4.000 0.7200 0.01552 0.00791 -0.0578 0.6525 1.0000
4.250 0.7434 0.01559 0.00792 -0.0563 0.6296 1.0000
4.500 0.7658 0.01570 0.00796 -0.0546 0.6042 1.0000
4.750 0.7870 0.01588 0.00808 -0.0529 0.5762 1.0000
5.000 0.8080 0.01611 0.00823 -0.0511 0.5462 1.0000
5.250 0.8286 0.01640 0.00839 -0.0493 0.5149 1.0000
5.500 0.8481 0.01679 0.00868 -0.0475 0.4817 1.0000
5.750 0.8675 0.01726 0.00901 -0.0457 0.4494 1.0000
6.000 0.8873 0.01782 0.00945 -0.0441 0.4200 1.0000
6.250 0.9074 0.01844 0.00992 -0.0426 0.3938 1.0000
6.500 0.9280 0.01911 0.01047 -0.0413 0.3703 1.0000
6.750 0.9488 0.01983 0.01112 -0.0400 0.3494 1.0000
7.000 0.9698 0.02057 0.01181 -0.0389 0.3303 1.0000
7.250 0.9914 0.02135 0.01256 -0.0378 0.3133 1.0000
7.500 1.0131 0.02215 0.01333 -0.0369 0.2976 1.0000
7.750 1.0346 0.02298 0.01416 -0.0359 0.2829 1.0000
8.000 1.0557 0.02383 0.01504 -0.0348 0.2686 1.0000
8.250 1.0762 0.02468 0.01594 -0.0337 0.2547 1.0000
8.500 1.0959 0.02553 0.01689 -0.0325 0.2410 1.0000
8.750 1.1146 0.02638 0.01780 -0.0312 0.2273 1.0000
9.000 1.1317 0.02716 0.01866 -0.0296 0.2133 1.0000
9.250 1.1460 0.02779 0.01933 -0.0277 0.1983 1.0000
9.500 1.1576 0.02832 0.01990 -0.0255 0.1830 1.0000
9.750 1.1674 0.02889 0.02055 -0.0230 0.1680 1.0000
10.000 1.1750 0.02956 0.02126 -0.0203 0.1529 1.0000
10.250 1.1792 0.03041 0.02220 -0.0172 0.1371 1.0000
10.500 1.1787 0.03144 0.02335 -0.0134 0.1201 1.0000
10.750 1.1755 0.03276 0.02466 -0.0096 0.1039 1.0000
11.000 1.1737 0.03441 0.02627 -0.0064 0.0901 1.0000
11.250 1.1742 0.03635 0.02822 -0.0037 0.0790 1.0000
11.500 1.1795 0.03857 0.03044 -0.0017 0.0715 1.0000
11.750 1.1862 0.04062 0.03254 0.0001 0.0657 1.0000
12.000 1.1970 0.04315 0.03511 0.0014 0.0612 1.0000
12.250 1.2003 0.04541 0.03767 0.0032 0.0576 1.0000
12.500 1.2062 0.04759 0.03992 0.0045 0.0547 1.0000
12.750 1.2217 0.05092 0.04324 0.0051 0.0519 1.0000
13.000 1.2179 0.05393 0.04659 0.0067 0.0510 1.0000
13.250 1.2105 0.05734 0.05034 0.0081 0.0502 1.0000
13.500 1.2005 0.06110 0.05441 0.0090 0.0497 1.0000
13.750 1.1870 0.06533 0.05893 0.0093 0.0495 1.0000
14.000 1.1700 0.07003 0.06391 0.0089 0.0494 1.0000
14.250 1.1495 0.07539 0.06954 0.0077 0.0494 1.0000
14.500 1.1255 0.08156 0.07595 0.0054 0.0495 1.0000
14.750 1.0992 0.08861 0.08323 0.0020 0.0499 1.0000
15.000 1.0714 0.09659 0.09141 -0.0025 0.0504 1.0000
15.250 1.0430 0.10551 0.10049 -0.0079 0.0511 1.0000
15.500 1.0119 0.11600 0.11110 -0.0144 0.0517 1.0000
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Polar data table (+)
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