CLARK Y AIRFOIL (clarky-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: CLARK Y AIRFOIL (clarky-il) Reynolds number: 500,000 Max Cl/Cd: 90.89 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-clarky-il-500000-n5.txt Download as CSV file: xf-clarky-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.8056 0.04280 0.03995 -0.0822 1.0000 0.0188
-12.250 -0.8431 0.03834 0.03530 -0.0811 1.0000 0.0188
-12.000 -0.8624 0.03532 0.03205 -0.0782 0.9996 0.0190
-11.750 -0.8498 0.03147 0.02773 -0.0810 0.9949 0.0196
-11.500 -0.8300 0.02925 0.02525 -0.0826 0.9910 0.0200
-11.250 -0.8037 0.02830 0.02424 -0.0838 0.9878 0.0204
-11.000 -0.7744 0.02770 0.02360 -0.0851 0.9853 0.0207
-10.750 -0.7444 0.02705 0.02290 -0.0866 0.9834 0.0212
-10.500 -0.7194 0.02620 0.02193 -0.0872 0.9789 0.0216
-10.250 -0.6939 0.02507 0.02064 -0.0880 0.9746 0.0222
-10.000 -0.6670 0.02386 0.01922 -0.0890 0.9712 0.0228
-9.750 -0.6427 0.02281 0.01797 -0.0892 0.9660 0.0234
-9.500 -0.6172 0.02189 0.01685 -0.0895 0.9606 0.0239
-9.250 -0.5903 0.02085 0.01561 -0.0901 0.9565 0.0243
-9.000 -0.5677 0.01980 0.01447 -0.0898 0.9498 0.0248
-8.750 -0.5414 0.01917 0.01378 -0.0899 0.9440 0.0253
-8.500 -0.5153 0.01858 0.01313 -0.0899 0.9383 0.0258
-8.250 -0.4904 0.01796 0.01241 -0.0896 0.9311 0.0262
-8.000 -0.4642 0.01737 0.01171 -0.0895 0.9252 0.0268
-7.750 -0.4394 0.01679 0.01102 -0.0891 0.9173 0.0273
-7.500 -0.4136 0.01616 0.01026 -0.0889 0.9105 0.0278
-7.250 -0.3885 0.01558 0.00957 -0.0884 0.9022 0.0282
-7.000 -0.3626 0.01504 0.00890 -0.0881 0.8949 0.0286
-6.750 -0.3370 0.01456 0.00833 -0.0877 0.8863 0.0289
-6.500 -0.3107 0.01413 0.00779 -0.0874 0.8787 0.0292
-6.250 -0.2850 0.01362 0.00719 -0.0871 0.8701 0.0297
-6.000 -0.2597 0.01300 0.00651 -0.0867 0.8624 0.0304
-5.750 -0.2336 0.01258 0.00604 -0.0864 0.8537 0.0309
-5.500 -0.2072 0.01221 0.00561 -0.0861 0.8459 0.0315
-5.250 -0.1807 0.01188 0.00522 -0.0858 0.8374 0.0320
-5.000 -0.1540 0.01158 0.00486 -0.0856 0.8295 0.0326
-4.750 -0.1272 0.01130 0.00452 -0.0853 0.8210 0.0332
-4.500 -0.1002 0.01105 0.00420 -0.0851 0.8129 0.0339
-4.250 -0.0733 0.01082 0.00390 -0.0848 0.8037 0.0346
-4.000 -0.0463 0.01062 0.00364 -0.0846 0.7932 0.0352
-3.750 -0.0194 0.01044 0.00338 -0.0843 0.7825 0.0360
-3.500 0.0074 0.01020 0.00309 -0.0840 0.7721 0.0376
-3.250 0.0346 0.01003 0.00288 -0.0838 0.7621 0.0390
-3.000 0.0617 0.00989 0.00268 -0.0836 0.7520 0.0405
-2.750 0.0889 0.00977 0.00251 -0.0834 0.7412 0.0423
-2.500 0.1162 0.00965 0.00235 -0.0832 0.7310 0.0456
-2.250 0.1433 0.00952 0.00222 -0.0829 0.7208 0.0527
-2.000 0.1702 0.00938 0.00212 -0.0827 0.7093 0.0714
-1.750 0.1972 0.00928 0.00202 -0.0825 0.6967 0.0875
-1.500 0.2242 0.00917 0.00193 -0.0823 0.6854 0.1051
-1.250 0.2512 0.00905 0.00186 -0.0821 0.6760 0.1316
-1.000 0.2781 0.00887 0.00181 -0.0819 0.6662 0.1766
-0.750 0.3050 0.00875 0.00177 -0.0818 0.6561 0.2148
-0.500 0.3318 0.00866 0.00175 -0.0816 0.6449 0.2547
-0.250 0.3585 0.00854 0.00174 -0.0814 0.6330 0.2998
0.000 0.3850 0.00839 0.00173 -0.0811 0.6216 0.3573
0.250 0.4113 0.00825 0.00174 -0.0809 0.6103 0.4205
0.500 0.4365 0.00802 0.00176 -0.0804 0.5985 0.5211
0.750 0.4588 0.00755 0.00183 -0.0793 0.5862 0.6956
1.000 0.4792 0.00724 0.00194 -0.0774 0.5730 0.8399
1.250 0.5069 0.00723 0.00204 -0.0769 0.5582 0.9260
1.500 0.5482 0.00736 0.00211 -0.0797 0.5414 0.9714
1.750 0.5852 0.00750 0.00216 -0.0816 0.5254 0.9868
2.000 0.6205 0.00766 0.00222 -0.0833 0.5086 0.9958
2.250 0.6517 0.00783 0.00229 -0.0842 0.4910 1.0000
2.500 0.6752 0.00800 0.00238 -0.0833 0.4723 1.0000
2.750 0.6985 0.00820 0.00247 -0.0824 0.4517 1.0000
3.000 0.7218 0.00842 0.00258 -0.0815 0.4312 1.0000
3.250 0.7451 0.00865 0.00271 -0.0806 0.4117 1.0000
3.500 0.7686 0.00888 0.00284 -0.0797 0.3925 1.0000
3.750 0.7923 0.00912 0.00299 -0.0790 0.3758 1.0000
4.000 0.8161 0.00937 0.00316 -0.0782 0.3604 1.0000
4.250 0.8397 0.00964 0.00334 -0.0775 0.3445 1.0000
4.500 0.8636 0.00991 0.00353 -0.0768 0.3306 1.0000
4.750 0.8876 0.01017 0.00373 -0.0761 0.3194 1.0000
5.000 0.9124 0.01038 0.00391 -0.0755 0.3096 1.0000
5.250 0.9370 0.01061 0.00410 -0.0750 0.3009 1.0000
5.500 0.9618 0.01082 0.00429 -0.0745 0.2929 1.0000
5.750 0.9867 0.01103 0.00450 -0.0740 0.2870 1.0000
6.000 1.0118 0.01123 0.00470 -0.0735 0.2805 1.0000
6.250 1.0362 0.01147 0.00493 -0.0730 0.2733 1.0000
6.500 1.0610 0.01168 0.00515 -0.0725 0.2647 1.0000
6.750 1.0852 0.01194 0.00539 -0.0719 0.2551 1.0000
7.000 1.1087 0.01224 0.00564 -0.0712 0.2411 1.0000
7.250 1.1312 0.01259 0.00593 -0.0705 0.2212 1.0000
7.500 1.1522 0.01307 0.00627 -0.0694 0.1962 1.0000
7.750 1.1711 0.01368 0.00672 -0.0682 0.1673 1.0000
8.000 1.1893 0.01434 0.00723 -0.0668 0.1425 1.0000
8.250 1.2079 0.01495 0.00775 -0.0654 0.1226 1.0000
8.500 1.2263 0.01555 0.00827 -0.0641 0.1069 1.0000
8.750 1.2447 0.01613 0.00879 -0.0627 0.0945 1.0000
9.000 1.2631 0.01664 0.00927 -0.0614 0.0862 1.0000
9.250 1.2801 0.01718 0.00980 -0.0597 0.0790 1.0000
9.500 1.2976 0.01768 0.01031 -0.0582 0.0727 1.0000
9.750 1.3142 0.01825 0.01088 -0.0567 0.0664 1.0000
10.000 1.3292 0.01893 0.01152 -0.0549 0.0579 1.0000
10.250 1.3418 0.01976 0.01228 -0.0529 0.0443 1.0000
10.500 1.3427 0.02137 0.01370 -0.0496 0.0187 1.0000
10.750 1.3533 0.02238 0.01473 -0.0476 0.0158 1.0000
11.000 1.3658 0.02327 0.01569 -0.0459 0.0144 1.0000
11.250 1.3779 0.02422 0.01670 -0.0442 0.0135 1.0000
11.500 1.3886 0.02530 0.01785 -0.0425 0.0127 1.0000
11.750 1.3978 0.02651 0.01915 -0.0408 0.0120 1.0000
12.000 1.4091 0.02761 0.02032 -0.0394 0.0116 1.0000
12.250 1.4190 0.02884 0.02164 -0.0380 0.0111 1.0000
12.500 1.4278 0.03019 0.02307 -0.0367 0.0106 1.0000
12.750 1.4352 0.03171 0.02467 -0.0353 0.0102 1.0000
13.000 1.4405 0.03347 0.02651 -0.0340 0.0098 1.0000
13.250 1.4428 0.03556 0.02870 -0.0327 0.0095 1.0000
13.500 1.4483 0.03742 0.03066 -0.0318 0.0093 1.0000
13.750 1.4528 0.03945 0.03278 -0.0309 0.0090 1.0000
14.000 1.4556 0.04169 0.03512 -0.0302 0.0088 1.0000
14.250 1.4571 0.04415 0.03769 -0.0296 0.0086 1.0000
14.500 1.4575 0.04682 0.04046 -0.0291 0.0084 1.0000
14.750 1.4566 0.04972 0.04346 -0.0288 0.0082 1.0000
15.000 1.4546 0.05285 0.04671 -0.0287 0.0081 1.0000
15.250 1.4516 0.05626 0.05023 -0.0289 0.0079 1.0000
15.500 1.4473 0.05992 0.05400 -0.0292 0.0078 1.0000
15.750 1.4415 0.06390 0.05809 -0.0298 0.0077 1.0000
16.000 1.4342 0.06821 0.06252 -0.0306 0.0076 1.0000
16.250 1.4250 0.07294 0.06736 -0.0317 0.0075 1.0000
16.500 1.4137 0.07807 0.07262 -0.0330 0.0074 1.0000
16.750 1.4010 0.08355 0.07821 -0.0346 0.0073 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CLARK Y AIRFOIL (clarky-il)