CLARK X AIRFOIL (clarkx-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: CLARK X AIRFOIL (clarkx-il) Reynolds number: 100,000 Max Cl/Cd: 55.22 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-clarkx-il-100000-n5.txt Download as CSV file: xf-clarkx-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: CLARK X 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4197 0.09161 0.08630 -0.0454 1.0000 0.0543
-9.500 -0.4281 0.08710 0.08187 -0.0469 1.0000 0.0544
-9.250 -0.4414 0.08248 0.07733 -0.0483 1.0000 0.0543
-9.000 -0.4623 0.07782 0.07278 -0.0491 1.0000 0.0542
-8.750 -0.4976 0.07291 0.06801 -0.0494 1.0000 0.0538
-8.500 -0.5371 0.06522 0.06034 -0.0517 1.0000 0.0533
-8.250 -0.5709 0.05815 0.05312 -0.0520 1.0000 0.0531
-8.000 -0.5631 0.05144 0.04602 -0.0574 0.9938 0.0546
-7.750 -0.5491 0.04516 0.03917 -0.0620 0.9864 0.0567
-7.500 -0.5316 0.03986 0.03319 -0.0649 0.9796 0.0581
-7.250 -0.5075 0.03583 0.02846 -0.0672 0.9741 0.0594
-7.000 -0.4831 0.03312 0.02520 -0.0684 0.9680 0.0609
-6.750 -0.4531 0.03168 0.02368 -0.0699 0.9632 0.0628
-6.500 -0.4214 0.03026 0.02204 -0.0716 0.9590 0.0648
-6.250 -0.3946 0.02875 0.02023 -0.0722 0.9524 0.0664
-6.000 -0.3616 0.02723 0.01840 -0.0738 0.9483 0.0684
-5.750 -0.3299 0.02592 0.01678 -0.0749 0.9436 0.0705
-5.500 -0.3006 0.02491 0.01547 -0.0754 0.9372 0.0732
-5.250 -0.2665 0.02391 0.01449 -0.0772 0.9332 0.0763
-5.000 -0.2338 0.02309 0.01359 -0.0783 0.9285 0.0794
-4.750 -0.2054 0.02233 0.01270 -0.0785 0.9216 0.0825
-4.500 -0.1710 0.02159 0.01176 -0.0798 0.9173 0.0862
-4.250 -0.1423 0.02092 0.01119 -0.0803 0.9110 0.0910
-4.000 -0.1124 0.02040 0.01061 -0.0807 0.9045 0.0970
-3.750 -0.0784 0.01974 0.00993 -0.0819 0.9003 0.1029
-3.500 -0.0529 0.01935 0.00955 -0.0816 0.8920 0.1102
-3.250 -0.0212 0.01885 0.00905 -0.0823 0.8866 0.1204
-3.000 0.0069 0.01844 0.00864 -0.0823 0.8797 0.1320
-2.750 0.0359 0.01798 0.00823 -0.0826 0.8729 0.1488
-2.500 0.0663 0.01748 0.00784 -0.0831 0.8672 0.1749
-2.250 0.0921 0.01707 0.00765 -0.0827 0.8589 0.2144
-2.000 0.1237 0.01661 0.00740 -0.0834 0.8539 0.2826
-1.750 0.1479 0.01634 0.00731 -0.0827 0.8446 0.3371
-1.500 0.1783 0.01587 0.00708 -0.0830 0.8390 0.4057
-1.250 0.2019 0.01550 0.00704 -0.0822 0.8297 0.4931
-1.000 0.2279 0.01485 0.00695 -0.0811 0.8238 0.6424
-0.750 0.2555 0.01445 0.00711 -0.0796 0.8153 0.8362
-0.500 0.3169 0.01424 0.00691 -0.0855 0.8102 0.9630
-0.250 0.3666 0.01417 0.00670 -0.0900 0.8022 1.0000
0.000 0.3932 0.01413 0.00652 -0.0895 0.7934 1.0000
0.250 0.4160 0.01419 0.00648 -0.0885 0.7824 1.0000
0.500 0.4415 0.01419 0.00637 -0.0878 0.7727 1.0000
0.750 0.4673 0.01420 0.00626 -0.0872 0.7626 1.0000
1.000 0.4909 0.01427 0.00627 -0.0862 0.7513 1.0000
1.250 0.5164 0.01432 0.00622 -0.0856 0.7410 1.0000
1.500 0.5426 0.01436 0.00617 -0.0850 0.7308 1.0000
1.750 0.5665 0.01446 0.00624 -0.0841 0.7191 1.0000
2.000 0.5915 0.01455 0.00627 -0.0834 0.7079 1.0000
2.250 0.6177 0.01462 0.00626 -0.0828 0.6972 1.0000
2.500 0.6424 0.01473 0.00634 -0.0821 0.6850 1.0000
2.750 0.6668 0.01485 0.00643 -0.0812 0.6721 1.0000
3.000 0.6916 0.01498 0.00653 -0.0805 0.6590 1.0000
3.250 0.7163 0.01510 0.00662 -0.0797 0.6453 1.0000
3.500 0.7408 0.01523 0.00671 -0.0788 0.6303 1.0000
3.750 0.7650 0.01537 0.00683 -0.0779 0.6143 1.0000
4.000 0.7891 0.01553 0.00695 -0.0770 0.5977 1.0000
4.250 0.8130 0.01571 0.00709 -0.0761 0.5810 1.0000
4.500 0.8368 0.01592 0.00728 -0.0752 0.5644 1.0000
4.750 0.8604 0.01615 0.00748 -0.0743 0.5478 1.0000
5.000 0.8839 0.01640 0.00771 -0.0734 0.5311 1.0000
5.250 0.9071 0.01667 0.00795 -0.0724 0.5141 1.0000
5.500 0.9299 0.01697 0.00825 -0.0714 0.4966 1.0000
5.750 0.9522 0.01730 0.00857 -0.0704 0.4782 1.0000
6.000 0.9743 0.01766 0.00892 -0.0693 0.4601 1.0000
6.250 0.9962 0.01804 0.00932 -0.0682 0.4426 1.0000
6.500 1.0177 0.01845 0.00972 -0.0671 0.4250 1.0000
6.750 1.0386 0.01890 0.01015 -0.0659 0.4074 1.0000
7.000 1.0592 0.01938 0.01061 -0.0647 0.3904 1.0000
7.250 1.0797 0.01988 0.01114 -0.0635 0.3746 1.0000
7.500 1.0999 0.02041 0.01169 -0.0622 0.3594 1.0000
7.750 1.1192 0.02098 0.01227 -0.0609 0.3436 1.0000
8.000 1.1367 0.02158 0.01287 -0.0593 0.3248 1.0000
8.250 1.1520 0.02222 0.01349 -0.0574 0.3031 1.0000
8.500 1.1652 0.02294 0.01415 -0.0553 0.2800 1.0000
8.750 1.1777 0.02368 0.01486 -0.0531 0.2560 1.0000
9.000 1.1890 0.02448 0.01560 -0.0509 0.2336 1.0000
9.250 1.1991 0.02531 0.01641 -0.0485 0.2122 1.0000
9.500 1.2084 0.02626 0.01732 -0.0461 0.1922 1.0000
9.750 1.2178 0.02730 0.01836 -0.0438 0.1692 1.0000
10.000 1.2244 0.02856 0.01956 -0.0414 0.1402 1.0000
10.250 1.2259 0.03028 0.02105 -0.0388 0.0997 1.0000
10.500 1.2229 0.03248 0.02298 -0.0360 0.0732 1.0000
10.750 1.2226 0.03459 0.02502 -0.0336 0.0590 1.0000
11.000 1.2224 0.03674 0.02717 -0.0315 0.0513 1.0000
11.250 1.2237 0.03884 0.02937 -0.0296 0.0468 1.0000
11.500 1.2231 0.04118 0.03179 -0.0280 0.0436 1.0000
11.750 1.2198 0.04386 0.03453 -0.0265 0.0412 1.0000
12.000 1.2198 0.04633 0.03720 -0.0254 0.0389 1.0000
12.250 1.2181 0.04906 0.04006 -0.0245 0.0370 1.0000
12.500 1.2147 0.05205 0.04317 -0.0238 0.0355 1.0000
12.750 1.2092 0.05538 0.04659 -0.0232 0.0344 1.0000
13.000 1.2041 0.05874 0.05005 -0.0228 0.0335 1.0000
13.250 1.2027 0.06178 0.05326 -0.0225 0.0326 1.0000
13.500 1.2014 0.06490 0.05652 -0.0222 0.0315 1.0000
13.750 1.2001 0.06808 0.05983 -0.0222 0.0305 1.0000
14.000 1.1987 0.07133 0.06319 -0.0222 0.0295 1.0000
14.250 1.1971 0.07464 0.06658 -0.0224 0.0286 1.0000
14.500 1.1962 0.07786 0.06984 -0.0225 0.0277 1.0000
14.750 1.1975 0.08074 0.07277 -0.0222 0.0269 1.0000
15.000 1.1992 0.08381 0.07602 -0.0222 0.0264 1.0000
15.250 1.2007 0.08697 0.07936 -0.0223 0.0258 1.0000
15.500 1.2009 0.09040 0.08296 -0.0226 0.0253 1.0000
15.750 1.1996 0.09414 0.08687 -0.0232 0.0248 1.0000
16.000 1.1964 0.09825 0.09116 -0.0243 0.0243 1.0000
16.250 1.1920 0.10266 0.09574 -0.0257 0.0238 1.0000
16.500 1.1868 0.10730 0.10055 -0.0274 0.0234 1.0000
16.750 1.1806 0.11222 0.10563 -0.0295 0.0229 1.0000
17.000 1.1742 0.11730 0.11086 -0.0318 0.0226 1.0000
17.250 1.1677 0.12247 0.11616 -0.0343 0.0223 1.0000
17.500 1.1615 0.12768 0.12149 -0.0369 0.0219 1.0000
17.750 1.1542 0.13329 0.12723 -0.0399 0.0217 1.0000
18.250 1.1326 0.14687 0.14113 -0.0480 0.0214 1.0000
18.500 1.1116 0.15712 0.15163 -0.0550 0.0216 1.0000
18.750 1.0022 0.20287 0.19777 -0.0830 0.0240 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CLARK X AIRFOIL (clarkx-il)