CLARK X AIRFOIL (clarkx-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: CLARK X AIRFOIL (clarkx-il) Reynolds number: 100,000 Max Cl/Cd: 54.2 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-clarkx-il-100000.txt Download as CSV file: xf-clarkx-il-100000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.3797 0.11686 0.11149 -0.0338 1.0000 0.1179
-10.250 -0.3719 0.11391 0.10854 -0.0338 1.0000 0.1220
-10.000 -0.3796 0.11142 0.10611 -0.0363 1.0000 0.1270
-9.750 -0.4196 0.11040 0.10529 -0.0428 1.0000 0.1289
-9.500 -0.3815 0.10430 0.09910 -0.0381 1.0000 0.1315
-9.250 -0.3678 0.10151 0.09631 -0.0365 1.0000 0.1357
-9.000 -0.3750 0.09913 0.09400 -0.0375 1.0000 0.1415
-8.750 -0.4174 0.09803 0.09311 -0.0400 1.0000 0.1439
-8.500 -0.3927 0.09332 0.08837 -0.0372 1.0000 0.1473
-8.250 -0.3838 0.09098 0.08605 -0.0351 1.0000 0.1514
-8.000 -0.3971 0.08914 0.08430 -0.0334 1.0000 0.1553
-7.750 -0.4337 0.08827 0.08361 -0.0305 1.0000 0.1579
-7.500 -0.4942 0.08664 0.08211 -0.0336 1.0000 0.1598
-7.250 -0.4858 0.08303 0.07857 -0.0294 1.0000 0.1618
-7.000 -0.4754 0.08115 0.07672 -0.0248 1.0000 0.1645
-6.750 -0.4805 0.07917 0.07478 -0.0226 1.0000 0.1683
-6.500 -0.5158 0.07493 0.07046 -0.0285 1.0000 0.1768
-6.250 -0.5066 0.07277 0.06839 -0.0241 1.0000 0.1789
-6.000 -0.5020 0.07106 0.06671 -0.0212 1.0000 0.1837
-5.750 -0.5080 0.06731 0.06287 -0.0236 1.0000 0.1941
-5.500 -0.5008 0.06566 0.06125 -0.0208 1.0000 0.1997
-5.250 -0.4788 0.04106 0.03446 -0.0402 0.9994 0.1121
-5.000 -0.4415 0.03714 0.02998 -0.0440 0.9950 0.1116
-4.750 -0.4036 0.03413 0.02636 -0.0471 0.9901 0.1124
-4.500 -0.3635 0.03172 0.02337 -0.0501 0.9851 0.1131
-4.250 -0.3257 0.02966 0.02100 -0.0529 0.9802 0.1164
-4.000 -0.2880 0.02857 0.01981 -0.0554 0.9741 0.1214
-3.750 -0.2454 0.02741 0.01827 -0.0584 0.9693 0.1261
-3.500 -0.2112 0.02630 0.01699 -0.0600 0.9622 0.1317
-3.250 -0.1683 0.02560 0.01620 -0.0631 0.9570 0.1411
-3.000 -0.1358 0.02476 0.01533 -0.0643 0.9495 0.1490
-2.750 -0.0950 0.02410 0.01466 -0.0670 0.9436 0.1630
-2.500 -0.0603 0.02350 0.01414 -0.0686 0.9366 0.1788
-2.250 -0.0233 0.02293 0.01375 -0.0706 0.9295 0.2046
-2.000 0.0152 0.02235 0.01348 -0.0728 0.9235 0.2529
-1.750 0.0465 0.02151 0.01329 -0.0738 0.9154 0.3634
-1.500 0.0853 0.02036 0.01321 -0.0757 0.9109 0.5817
-1.250 0.1177 0.01972 0.01355 -0.0742 0.9037 0.9396
-1.000 0.1745 0.01971 0.01325 -0.0800 0.8977 1.0000
-0.750 0.2067 0.01985 0.01319 -0.0810 0.8883 1.0000
-0.500 0.2514 0.01981 0.01296 -0.0842 0.8817 1.0000
-0.250 0.2810 0.01994 0.01295 -0.0846 0.8715 1.0000
0.000 0.3266 0.01975 0.01261 -0.0876 0.8655 1.0000
0.250 0.3528 0.01987 0.01264 -0.0872 0.8544 1.0000
0.500 0.3980 0.01953 0.01220 -0.0898 0.8489 1.0000
1.000 0.4575 0.01946 0.01199 -0.0898 0.8293 1.0000
1.250 0.4883 0.01935 0.01183 -0.0898 0.8196 1.0000
1.500 0.5149 0.01938 0.01181 -0.0892 0.8089 1.0000
1.750 0.5529 0.01899 0.01138 -0.0902 0.8020 1.0000
2.000 0.5764 0.01908 0.01144 -0.0891 0.7897 1.0000
2.250 0.6046 0.01901 0.01135 -0.0885 0.7792 1.0000
2.500 0.6394 0.01864 0.01095 -0.0888 0.7707 1.0000
2.750 0.6636 0.01868 0.01098 -0.0877 0.7577 1.0000
3.000 0.6901 0.01864 0.01093 -0.0869 0.7453 1.0000
3.250 0.7191 0.01846 0.01074 -0.0863 0.7334 1.0000
3.500 0.7506 0.01816 0.01039 -0.0860 0.7217 1.0000
3.750 0.7779 0.01803 0.01023 -0.0851 0.7075 1.0000
4.000 0.8037 0.01798 0.01018 -0.0841 0.6924 1.0000
4.250 0.8295 0.01796 0.01015 -0.0831 0.6771 1.0000
4.500 0.8554 0.01797 0.01015 -0.0822 0.6615 1.0000
4.750 0.8813 0.01799 0.01016 -0.0812 0.6455 1.0000
5.000 0.9073 0.01803 0.01018 -0.0804 0.6290 1.0000
5.250 0.9308 0.01819 0.01034 -0.0791 0.6107 1.0000
5.500 0.9542 0.01837 0.01054 -0.0780 0.5920 1.0000
5.750 0.9783 0.01854 0.01071 -0.0769 0.5728 1.0000
6.000 1.0030 0.01873 0.01085 -0.0758 0.5534 1.0000
6.250 1.0268 0.01901 0.01108 -0.0748 0.5336 1.0000
6.500 1.0484 0.01939 0.01148 -0.0734 0.5126 1.0000
6.750 1.0715 0.01977 0.01183 -0.0723 0.4923 1.0000
7.000 1.0949 0.02021 0.01218 -0.0712 0.4725 1.0000
7.250 1.1151 0.02075 0.01277 -0.0698 0.4517 1.0000
7.500 1.1362 0.02127 0.01326 -0.0685 0.4313 1.0000
7.750 1.1556 0.02171 0.01361 -0.0668 0.4094 1.0000
8.000 1.1709 0.02212 0.01402 -0.0646 0.3850 1.0000
8.250 1.1862 0.02258 0.01441 -0.0624 0.3620 1.0000
8.500 1.2000 0.02308 0.01493 -0.0601 0.3396 1.0000
8.750 1.2142 0.02365 0.01544 -0.0579 0.3193 1.0000
9.000 1.2272 0.02425 0.01612 -0.0556 0.2997 1.0000
9.250 1.2382 0.02486 0.01681 -0.0531 0.2792 1.0000
9.500 1.2459 0.02556 0.01751 -0.0501 0.2573 1.0000
9.750 1.2474 0.02636 0.01836 -0.0463 0.2292 1.0000
10.000 1.2414 0.02772 0.01960 -0.0418 0.1860 1.0000
10.250 1.2273 0.03023 0.02165 -0.0372 0.1285 1.0000
10.500 1.2184 0.03284 0.02393 -0.0336 0.1013 1.0000
10.750 1.2136 0.03530 0.02624 -0.0307 0.0894 1.0000
11.000 1.2134 0.03750 0.02848 -0.0284 0.0811 1.0000
11.250 1.2133 0.03983 0.03079 -0.0262 0.0756 1.0000
11.500 1.2176 0.04188 0.03291 -0.0243 0.0710 1.0000
11.750 1.2239 0.04404 0.03494 -0.0224 0.0674 1.0000
12.000 1.2359 0.04584 0.03688 -0.0209 0.0641 1.0000
12.250 1.2474 0.04769 0.03879 -0.0195 0.0608 1.0000
12.500 1.2672 0.04948 0.04044 -0.0184 0.0575 1.0000
12.750 1.2916 0.05163 0.04271 -0.0174 0.0556 1.0000
13.000 1.3056 0.05397 0.04533 -0.0162 0.0543 1.0000
13.250 1.3151 0.05655 0.04815 -0.0150 0.0530 1.0000
13.500 1.3215 0.05921 0.05102 -0.0137 0.0516 1.0000
13.750 1.3264 0.06194 0.05392 -0.0126 0.0503 1.0000
14.000 1.3296 0.06495 0.05709 -0.0116 0.0493 1.0000
14.250 1.3306 0.06841 0.06074 -0.0107 0.0487 1.0000
14.500 1.3245 0.07227 0.06485 -0.0098 0.0485 1.0000
14.750 1.3125 0.07645 0.06932 -0.0091 0.0485 1.0000
15.000 1.2953 0.08106 0.07422 -0.0090 0.0487 1.0000
15.250 1.2728 0.08633 0.07980 -0.0097 0.0490 1.0000
15.500 1.2430 0.09269 0.08650 -0.0115 0.0495 1.0000
15.750 1.1991 0.10148 0.09568 -0.0158 0.0503 1.0000
16.000 1.1533 0.11236 0.10691 -0.0224 0.0514 1.0000
16.250 1.1039 0.12608 0.12091 -0.0316 0.0529 1.0000
16.500 1.0566 0.14186 0.13683 -0.0420 0.0550 1.0000
16.750 1.0327 0.15340 0.14839 -0.0486 0.0563 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CLARK X AIRFOIL (clarkx-il)