CLARK K AIRFOIL (clarkk-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: CLARK K AIRFOIL (clarkk-il) Reynolds number: 50,000 Max Cl/Cd: 32.18 at α=10° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-clarkk-il-50000.txt Download as CSV file: xf-clarkk-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: CLARK K 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.3670 0.12687 0.11916 -0.0249 1.0000 0.2386
-10.250 -0.3559 0.12345 0.11575 -0.0247 1.0000 0.2465
-10.000 -0.3858 0.12517 0.11762 -0.0266 1.0000 0.2522
-9.750 -0.3481 0.11753 0.10989 -0.0252 1.0000 0.2592
-9.500 -0.3609 0.11696 0.10942 -0.0258 1.0000 0.2683
-9.250 -0.3427 0.11208 0.10454 -0.0251 1.0000 0.2750
-9.000 -0.3520 0.11103 0.10358 -0.0251 1.0000 0.2850
-8.750 -0.3400 0.10689 0.09947 -0.0246 1.0000 0.2910
-8.500 -0.3420 0.10503 0.09768 -0.0239 1.0000 0.3010
-8.250 -0.3409 0.10205 0.09478 -0.0232 1.0000 0.3072
-8.000 -0.3404 0.09997 0.09277 -0.0218 1.0000 0.3177
-7.750 -0.3384 0.09711 0.08998 -0.0205 1.0000 0.3255
-7.500 -0.3603 0.09691 0.08994 -0.0180 1.0000 0.3368
-7.250 -0.3388 0.09274 0.08576 -0.0167 1.0000 0.3478
-7.000 -0.3695 0.09250 0.08572 -0.0130 1.0000 0.3565
-6.750 -0.3667 0.09032 0.08360 -0.0103 1.0000 0.3699
-6.500 -0.3596 0.08734 0.08066 -0.0081 1.0000 0.3796
-6.250 -0.4217 0.08942 0.08303 -0.0014 1.0000 0.3892
-6.000 -0.3931 0.08519 0.07877 0.0004 1.0000 0.4031
-5.750 -0.3935 0.08287 0.07652 0.0035 1.0000 0.4150
-5.500 -0.4060 0.08131 0.07507 0.0073 1.0000 0.4293
-5.250 -0.4145 0.07961 0.07345 0.0110 1.0000 0.4459
-5.000 -0.4161 0.07770 0.07161 0.0146 1.0000 0.4655
-4.750 -0.4435 0.07731 0.07134 0.0195 1.0000 0.4918
-4.500 -0.4175 0.07406 0.06810 0.0230 1.0000 0.5185
-4.250 -0.4145 0.05292 0.04529 -0.0320 1.0000 0.2188
-4.000 -0.3918 0.04846 0.04047 -0.0342 1.0000 0.2036
-3.750 -0.3688 0.04540 0.03712 -0.0355 1.0000 0.1948
-3.500 -0.3390 0.04164 0.03249 -0.0383 1.0000 0.1837
-3.250 -0.3158 0.03957 0.03017 -0.0388 1.0000 0.1811
-3.000 -0.2917 0.03782 0.02808 -0.0394 1.0000 0.1804
-2.750 -0.2676 0.03641 0.02634 -0.0398 1.0000 0.1809
-2.500 -0.2436 0.03519 0.02482 -0.0400 1.0000 0.1811
-2.250 -0.2196 0.03414 0.02349 -0.0402 1.0000 0.1809
-2.000 -0.1959 0.03332 0.02242 -0.0402 1.0000 0.1813
-1.750 -0.1726 0.03270 0.02158 -0.0402 1.0000 0.1825
-1.500 -0.1497 0.03227 0.02093 -0.0400 1.0000 0.1845
-1.250 -0.1282 0.03193 0.02063 -0.0399 1.0000 0.1887
-1.000 -0.1069 0.03183 0.02051 -0.0396 1.0000 0.1948
-0.750 -0.0847 0.03186 0.02038 -0.0394 1.0000 0.2012
-0.500 -0.0620 0.03186 0.02046 -0.0395 1.0000 0.2085
-0.250 -0.0116 0.03217 0.02081 -0.0443 0.9903 0.2267
0.000 0.0400 0.03216 0.02120 -0.0495 0.9799 0.2812
0.500 0.1189 0.03117 0.02221 -0.0539 0.9566 1.0000
0.750 0.1563 0.03214 0.02290 -0.0567 0.9432 1.0000
1.000 0.1924 0.03311 0.02364 -0.0592 0.9295 1.0000
1.250 0.2282 0.03407 0.02441 -0.0616 0.9158 1.0000
1.500 0.2643 0.03502 0.02520 -0.0640 0.9019 1.0000
1.750 0.3013 0.03594 0.02598 -0.0663 0.8877 1.0000
2.000 0.3395 0.03681 0.02674 -0.0688 0.8736 1.0000
2.250 0.3801 0.03760 0.02743 -0.0714 0.8592 1.0000
2.500 0.4132 0.03832 0.02809 -0.0727 0.8438 1.0000
2.750 0.4452 0.03900 0.02873 -0.0737 0.8279 1.0000
3.000 0.4783 0.03961 0.02930 -0.0747 0.8118 1.0000
3.250 0.5123 0.04012 0.02980 -0.0756 0.7954 1.0000
3.500 0.5470 0.04051 0.03019 -0.0764 0.7789 1.0000
3.750 0.5819 0.04079 0.03049 -0.0770 0.7624 1.0000
4.000 0.6165 0.04098 0.03072 -0.0774 0.7459 1.0000
4.250 0.6517 0.04102 0.03080 -0.0776 0.7294 1.0000
4.500 0.6868 0.04095 0.03080 -0.0776 0.7131 1.0000
4.750 0.7219 0.04076 0.03068 -0.0774 0.6968 1.0000
5.000 0.7572 0.04040 0.03040 -0.0770 0.6806 1.0000
5.250 0.7913 0.04003 0.03012 -0.0763 0.6643 1.0000
5.500 0.8270 0.03945 0.02963 -0.0756 0.6480 1.0000
5.750 0.8630 0.03881 0.02910 -0.0748 0.6316 1.0000
6.000 0.8996 0.03812 0.02849 -0.0741 0.6150 1.0000
6.250 0.9207 0.03853 0.02899 -0.0723 0.5958 1.0000
6.500 0.9463 0.03872 0.02927 -0.0709 0.5776 1.0000
6.750 0.9754 0.03870 0.02934 -0.0697 0.5600 1.0000
7.000 1.0041 0.03880 0.02952 -0.0686 0.5430 1.0000
7.250 1.0315 0.03908 0.02991 -0.0676 0.5264 1.0000
7.500 1.0573 0.03954 0.03046 -0.0664 0.5102 1.0000
7.750 1.0815 0.04023 0.03124 -0.0653 0.4949 1.0000
8.000 1.1059 0.04091 0.03204 -0.0641 0.4799 1.0000
8.250 1.1319 0.04155 0.03277 -0.0631 0.4654 1.0000
8.500 1.1583 0.04209 0.03340 -0.0621 0.4505 1.0000
8.750 1.1825 0.04269 0.03412 -0.0608 0.4348 1.0000
9.000 1.1996 0.04335 0.03491 -0.0586 0.4168 1.0000
9.250 1.2383 0.04074 0.03209 -0.0568 0.3856 1.0000
9.500 1.2618 0.03926 0.03041 -0.0542 0.3550 1.0000
9.750 1.2693 0.03983 0.03109 -0.0509 0.3323 1.0000
10.000 1.2834 0.03988 0.03103 -0.0481 0.3057 1.0000
10.250 1.2877 0.04034 0.03136 -0.0442 0.2753 1.0000
10.500 1.2810 0.04157 0.03247 -0.0393 0.2422 1.0000
10.750 1.2702 0.04334 0.03402 -0.0342 0.2098 1.0000
11.000 1.2655 0.04532 0.03568 -0.0301 0.1832 1.0000
11.250 1.2628 0.04775 0.03815 -0.0268 0.1649 1.0000
11.500 1.2666 0.05005 0.04042 -0.0244 0.1508 1.0000
11.750 1.2765 0.05232 0.04261 -0.0227 0.1394 1.0000
12.000 1.2849 0.05491 0.04534 -0.0209 0.1313 1.0000
12.250 1.2924 0.05794 0.04851 -0.0194 0.1251 1.0000
12.500 1.2904 0.06108 0.05191 -0.0174 0.1206 1.0000
12.750 0.8713 0.10054 0.09351 -0.0206 0.2027 1.0000
13.000 0.8264 0.11250 0.10541 -0.0247 0.2009 1.0000
|
Polar data table (+)
Polar graphs
<< Back to CLARK K AIRFOIL (clarkk-il)