CLARK Y AIRFOIL (clarky-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: CLARK Y AIRFOIL (clarky-il) Reynolds number: 50,000 Max Cl/Cd: 36.54 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-clarky-il-50000-n5.txt Download as CSV file: xf-clarky-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3683 0.12749 0.11999 -0.0321 1.0000 0.1294
-10.500 -0.3799 0.12552 0.11811 -0.0349 1.0000 0.1311
-10.250 -0.3871 0.12285 0.11552 -0.0371 1.0000 0.1315
-10.000 -0.3561 0.11742 0.11002 -0.0338 1.0000 0.1362
-9.750 -0.3865 0.10888 0.10159 -0.0427 1.0000 0.0878
-9.500 -0.3694 0.10555 0.09820 -0.0406 1.0000 0.0851
-9.250 -0.3674 0.10205 0.09476 -0.0408 1.0000 0.0829
-9.000 -0.3705 0.09846 0.09122 -0.0415 1.0000 0.0808
-8.750 -0.3776 0.09491 0.08777 -0.0421 1.0000 0.0795
-8.250 -0.4394 0.08464 0.07783 -0.0474 1.0000 0.0740
-8.000 -0.4506 0.08108 0.07432 -0.0472 1.0000 0.0739
-7.750 -0.4626 0.07740 0.07067 -0.0469 1.0000 0.0738
-7.500 -0.4743 0.07361 0.06686 -0.0465 1.0000 0.0738
-7.250 -0.4843 0.06974 0.06291 -0.0461 1.0000 0.0738
-7.000 -0.4913 0.06583 0.05888 -0.0457 1.0000 0.0738
-6.750 -0.4944 0.06200 0.05487 -0.0453 1.0000 0.0737
-6.500 -0.4936 0.05827 0.05092 -0.0448 1.0000 0.0736
-6.250 -0.4896 0.05457 0.04694 -0.0445 1.0000 0.0735
-6.000 -0.4821 0.05100 0.04301 -0.0443 1.0000 0.0736
-5.750 -0.4540 0.04675 0.03823 -0.0478 0.9942 0.0746
-5.500 -0.4248 0.04466 0.03605 -0.0499 0.9882 0.0770
-5.250 -0.3926 0.04209 0.03312 -0.0527 0.9827 0.0793
-5.000 -0.3627 0.03934 0.02988 -0.0547 0.9761 0.0807
-4.750 -0.3280 0.03687 0.02688 -0.0571 0.9709 0.0823
-4.500 -0.2966 0.03491 0.02444 -0.0586 0.9642 0.0845
-4.250 -0.2616 0.03329 0.02236 -0.0605 0.9584 0.0884
-4.000 -0.2293 0.03213 0.02111 -0.0619 0.9522 0.0922
-3.750 -0.1962 0.03099 0.01975 -0.0632 0.9456 0.0959
-3.500 -0.1577 0.02990 0.01834 -0.0652 0.9408 0.1008
-3.250 -0.1300 0.02912 0.01758 -0.0655 0.9325 0.1071
-3.000 -0.0927 0.02836 0.01668 -0.0674 0.9270 0.1181
-2.750 -0.0630 0.02772 0.01598 -0.0679 0.9193 0.1302
-2.500 -0.0276 0.02702 0.01537 -0.0696 0.9130 0.1521
-2.250 0.0039 0.02644 0.01494 -0.0707 0.9057 0.1918
-2.000 0.0363 0.02580 0.01472 -0.0721 0.8985 0.2686
-1.750 0.0685 0.02503 0.01447 -0.0733 0.8921 0.3781
-1.500 0.0939 0.02427 0.01439 -0.0728 0.8840 0.5183
-1.250 0.1299 0.02332 0.01447 -0.0719 0.8799 0.8331
-1.000 0.1803 0.02334 0.01424 -0.0763 0.8714 1.0000
-0.750 0.2176 0.02339 0.01398 -0.0781 0.8645 1.0000
-0.500 0.2416 0.02361 0.01395 -0.0777 0.8536 1.0000
-0.250 0.2742 0.02368 0.01379 -0.0785 0.8446 1.0000
0.000 0.3084 0.02361 0.01351 -0.0793 0.8341 1.0000
0.250 0.3367 0.02360 0.01332 -0.0790 0.8210 1.0000
0.500 0.3660 0.02355 0.01311 -0.0789 0.8082 1.0000
0.750 0.3972 0.02346 0.01288 -0.0789 0.7967 1.0000
1.000 0.4301 0.02334 0.01263 -0.0793 0.7868 1.0000
1.250 0.4543 0.02348 0.01268 -0.0785 0.7743 1.0000
1.500 0.4808 0.02357 0.01268 -0.0779 0.7626 1.0000
1.750 0.5119 0.02350 0.01252 -0.0780 0.7527 1.0000
2.000 0.5391 0.02355 0.01250 -0.0775 0.7408 1.0000
2.250 0.5637 0.02369 0.01260 -0.0766 0.7279 1.0000
2.500 0.5897 0.02377 0.01264 -0.0759 0.7152 1.0000
2.750 0.6171 0.02381 0.01263 -0.0754 0.7027 1.0000
3.000 0.6462 0.02377 0.01256 -0.0750 0.6908 1.0000
3.250 0.6728 0.02383 0.01260 -0.0743 0.6773 1.0000
3.500 0.6976 0.02396 0.01272 -0.0734 0.6626 1.0000
3.750 0.7225 0.02410 0.01286 -0.0725 0.6478 1.0000
4.000 0.7476 0.02425 0.01300 -0.0717 0.6327 1.0000
4.250 0.7728 0.02441 0.01315 -0.0708 0.6173 1.0000
4.500 0.7979 0.02459 0.01334 -0.0700 0.6017 1.0000
4.750 0.8229 0.02481 0.01355 -0.0691 0.5861 1.0000
5.000 0.8478 0.02506 0.01380 -0.0683 0.5704 1.0000
5.250 0.8724 0.02535 0.01411 -0.0675 0.5549 1.0000
5.500 0.8967 0.02571 0.01447 -0.0667 0.5398 1.0000
5.750 0.9206 0.02611 0.01489 -0.0658 0.5250 1.0000
6.000 0.9442 0.02655 0.01537 -0.0650 0.5105 1.0000
6.250 0.9676 0.02703 0.01589 -0.0642 0.4965 1.0000
6.500 0.9908 0.02753 0.01642 -0.0633 0.4828 1.0000
6.750 1.0137 0.02803 0.01695 -0.0624 0.4694 1.0000
7.000 1.0365 0.02854 0.01751 -0.0615 0.4560 1.0000
7.250 1.0589 0.02904 0.01803 -0.0605 0.4426 1.0000
7.500 1.0802 0.02956 0.01859 -0.0593 0.4286 1.0000
7.750 1.0990 0.03014 0.01923 -0.0579 0.4135 1.0000
8.000 1.1171 0.03072 0.01985 -0.0563 0.3981 1.0000
8.250 1.1344 0.03136 0.02052 -0.0547 0.3829 1.0000
8.500 1.1507 0.03207 0.02129 -0.0530 0.3680 1.0000
8.750 1.1663 0.03284 0.02218 -0.0513 0.3538 1.0000
9.000 1.1812 0.03368 0.02312 -0.0496 0.3398 1.0000
9.250 1.1953 0.03458 0.02412 -0.0478 0.3261 1.0000
9.500 1.2082 0.03551 0.02516 -0.0459 0.3122 1.0000
9.750 1.2198 0.03650 0.02626 -0.0438 0.2985 1.0000
10.000 1.2292 0.03753 0.02741 -0.0415 0.2851 1.0000
10.250 1.2370 0.03863 0.02860 -0.0392 0.2718 1.0000
10.500 1.2441 0.03980 0.02984 -0.0368 0.2588 1.0000
10.750 1.2481 0.04125 0.03147 -0.0345 0.2459 1.0000
11.000 1.2524 0.04284 0.03320 -0.0324 0.2338 1.0000
11.250 1.2570 0.04447 0.03496 -0.0304 0.2226 1.0000
11.500 1.2624 0.04606 0.03662 -0.0286 0.2125 1.0000
11.750 1.2666 0.04784 0.03850 -0.0269 0.2028 1.0000
12.000 1.2690 0.04994 0.04076 -0.0254 0.1939 1.0000
12.250 1.2747 0.05164 0.04246 -0.0240 0.1859 1.0000
12.500 1.2723 0.05423 0.04524 -0.0227 0.1775 1.0000
12.750 1.2707 0.05664 0.04770 -0.0216 0.1694 1.0000
13.000 1.2669 0.05930 0.05044 -0.0208 0.1611 1.0000
13.250 1.2598 0.06261 0.05389 -0.0203 0.1535 1.0000
13.500 1.2563 0.06533 0.05660 -0.0198 0.1456 1.0000
13.750 1.2448 0.06962 0.06115 -0.0201 0.1389 1.0000
14.000 1.2399 0.07271 0.06416 -0.0202 0.1313 1.0000
14.250 1.2245 0.07794 0.06970 -0.0214 0.1250 1.0000
14.500 1.2176 0.08161 0.07330 -0.0221 0.1177 1.0000
14.750 1.2014 0.08752 0.07952 -0.0240 0.1120 1.0000
15.000 1.1960 0.09127 0.08321 -0.0250 0.1052 1.0000
15.250 1.1795 0.09779 0.09003 -0.0275 0.1005 1.0000
15.500 1.1711 0.10264 0.09495 -0.0292 0.0944 1.0000
15.750 1.1586 0.10858 0.10103 -0.0316 0.0893 1.0000
16.000 1.1435 0.11533 0.10796 -0.0346 0.0844 1.0000
16.250 1.1380 0.11991 0.11252 -0.0365 0.0785 1.0000
16.500 1.1148 0.12918 0.12205 -0.0412 0.0752 1.0000
16.750 1.1211 0.13104 0.12374 -0.0416 0.0685 1.0000
17.000 1.0920 0.14257 0.13553 -0.0479 0.0668 1.0000
17.250 1.0525 0.15811 0.15118 -0.0568 0.0657 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CLARK Y AIRFOIL (clarky-il)