DAE-21 AIRFOIL (dae21-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: DAE-21 AIRFOIL (dae21-il) Reynolds number: 200,000 Max Cl/Cd: 86.26 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-dae21-il-200000.txt Download as CSV file: xf-dae21-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: DAE-21 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.0228 0.09547 0.09143 -0.0829 0.8254 0.0370
-8.500 -0.0169 0.09313 0.08910 -0.0844 0.8217 0.0386
-8.250 -0.0215 0.09159 0.08757 -0.0878 0.8179 0.0395
-8.000 -0.0211 0.08902 0.08502 -0.0901 0.8147 0.0399
-7.750 0.0001 0.08565 0.08165 -0.0885 0.8120 0.0407
-7.500 0.0104 0.08331 0.07933 -0.0890 0.8081 0.0416
-7.250 0.0181 0.08105 0.07708 -0.0899 0.8045 0.0427
-7.000 0.0236 0.07880 0.07484 -0.0908 0.8012 0.0438
-6.750 0.0242 0.07689 0.07293 -0.0921 0.7981 0.0451
-6.500 0.0237 0.07447 0.07058 -0.1022 0.7920 0.0462
-6.250 0.0337 0.07112 0.06725 -0.1022 0.7886 0.0467
-6.000 0.0464 0.06896 0.06508 -0.1001 0.7859 0.0476
-5.750 0.0592 0.06673 0.06283 -0.1008 0.7833 0.0487
-5.500 0.0731 0.06425 0.06033 -0.1036 0.7800 0.0506
-5.250 0.1020 0.05814 0.05404 -0.1210 0.7741 0.0543
-5.000 0.1127 0.05632 0.05231 -0.1178 0.7710 0.0551
-4.750 0.1284 0.05448 0.05046 -0.1173 0.7682 0.0566
-4.500 0.1667 0.04922 0.04465 -0.1293 0.7655 0.0635
-4.250 0.1806 0.04684 0.04245 -0.1280 0.7620 0.0646
-4.000 0.1986 0.04559 0.04128 -0.1278 0.7574 0.0669
-3.750 0.2305 0.04182 0.03714 -0.1331 0.7539 0.0757
-3.500 0.2509 0.04025 0.03562 -0.1326 0.7511 0.0785
-3.250 0.2825 0.03735 0.03227 -0.1357 0.7489 0.0886
-3.000 0.3052 0.03576 0.03072 -0.1357 0.7463 0.0920
-2.750 0.3322 0.03410 0.02881 -0.1376 0.7406 0.1037
-2.500 0.3571 0.03271 0.02741 -0.1377 0.7370 0.1095
-2.250 0.4020 0.02556 0.01862 -0.1389 0.7348 0.0531
-2.000 0.4310 0.02413 0.01687 -0.1388 0.7325 0.0530
-1.750 0.4601 0.02276 0.01522 -0.1385 0.7306 0.0524
-1.500 0.4849 0.02231 0.01465 -0.1384 0.7247 0.0527
-1.250 0.5123 0.02179 0.01396 -0.1381 0.7208 0.0540
-1.000 0.5401 0.02092 0.01314 -0.1380 0.7180 0.0583
-0.750 0.5689 0.02030 0.01244 -0.1377 0.7157 0.0619
-0.500 0.5977 0.01959 0.01174 -0.1374 0.7138 0.0673
-0.250 0.6204 0.01974 0.01199 -0.1372 0.7072 0.0772
0.000 0.6481 0.01932 0.01165 -0.1371 0.7034 0.1020
0.250 0.6738 0.01785 0.01150 -0.1371 0.7007 0.4969
0.750 0.7330 0.01664 0.01100 -0.1364 0.6961 1.0000
1.000 0.7556 0.01720 0.01152 -0.1361 0.6890 1.0000
1.250 0.7841 0.01718 0.01136 -0.1360 0.6856 1.0000
1.500 0.8139 0.01704 0.01109 -0.1359 0.6831 1.0000
1.750 0.8443 0.01689 0.01080 -0.1359 0.6810 1.0000
2.000 0.8661 0.01746 0.01138 -0.1355 0.6738 1.0000
2.250 0.8942 0.01745 0.01130 -0.1353 0.6699 1.0000
2.500 0.9243 0.01726 0.01101 -0.1353 0.6671 1.0000
2.750 0.9553 0.01705 0.01068 -0.1354 0.6649 1.0000
3.000 0.9765 0.01757 0.01126 -0.1348 0.6573 1.0000
3.250 1.0052 0.01747 0.01111 -0.1347 0.6533 1.0000
3.500 1.0361 0.01721 0.01076 -0.1348 0.6505 1.0000
3.750 1.0622 0.01734 0.01088 -0.1345 0.6454 1.0000
4.000 1.0874 0.01747 0.01103 -0.1341 0.6393 1.0000
4.250 1.1178 0.01721 0.01072 -0.1342 0.6356 1.0000
4.500 1.1486 0.01699 0.01042 -0.1343 0.6322 1.0000
4.750 1.1701 0.01732 0.01085 -0.1336 0.6242 1.0000
5.000 1.2003 0.01707 0.01055 -0.1336 0.6198 1.0000
5.250 1.2276 0.01706 0.01053 -0.1334 0.6144 1.0000
5.500 1.2524 0.01713 0.01064 -0.1329 0.6072 1.0000
5.750 1.2836 0.01685 0.01031 -0.1331 0.6026 1.0000
6.000 1.3057 0.01710 0.01064 -0.1324 0.5945 1.0000
6.250 1.3346 0.01693 0.01044 -0.1323 0.5884 1.0000
6.500 1.3585 0.01706 0.01064 -0.1317 0.5805 1.0000
6.750 1.3858 0.01698 0.01054 -0.1315 0.5731 1.0000
7.000 1.4090 0.01711 0.01073 -0.1307 0.5643 1.0000
7.250 1.4363 0.01703 0.01062 -0.1305 0.5561 1.0000
7.500 1.4571 0.01725 0.01092 -0.1295 0.5457 1.0000
7.750 1.4818 0.01731 0.01097 -0.1289 0.5362 1.0000
8.000 1.5044 0.01744 0.01111 -0.1280 0.5254 1.0000
8.250 1.5232 0.01772 0.01147 -0.1267 0.5131 1.0000
8.500 1.5427 0.01798 0.01175 -0.1254 0.5008 1.0000
8.750 1.5615 0.01828 0.01204 -0.1241 0.4879 1.0000
9.000 1.5786 0.01864 0.01240 -0.1225 0.4744 1.0000
9.250 1.5933 0.01909 0.01283 -0.1206 0.4603 1.0000
9.500 1.6044 0.01964 0.01337 -0.1182 0.4458 1.0000
9.750 1.6096 0.02031 0.01405 -0.1149 0.4314 1.0000
10.000 1.6137 0.02113 0.01488 -0.1118 0.4168 1.0000
10.250 1.6172 0.02211 0.01585 -0.1089 0.4018 1.0000
10.500 1.6203 0.02322 0.01695 -0.1061 0.3865 1.0000
10.750 1.6226 0.02449 0.01819 -0.1035 0.3707 1.0000
11.000 1.6242 0.02589 0.01956 -0.1010 0.3548 1.0000
11.250 1.6253 0.02742 0.02108 -0.0986 0.3389 1.0000
11.500 1.6257 0.02909 0.02271 -0.0963 0.3230 1.0000
11.750 1.6255 0.03089 0.02448 -0.0941 0.3071 1.0000
12.000 1.6247 0.03280 0.02637 -0.0920 0.2915 1.0000
12.250 1.6235 0.03484 0.02837 -0.0901 0.2761 1.0000
12.500 1.6217 0.03701 0.03052 -0.0882 0.2608 1.0000
12.750 1.6195 0.03929 0.03279 -0.0865 0.2459 1.0000
13.000 1.6173 0.04169 0.03517 -0.0849 0.2313 1.0000
13.250 1.6146 0.04425 0.03772 -0.0835 0.2167 1.0000
13.500 1.6122 0.04692 0.04037 -0.0823 0.2027 1.0000
13.750 1.6092 0.04973 0.04317 -0.0812 0.1889 1.0000
14.000 1.6062 0.05267 0.04610 -0.0802 0.1758 1.0000
14.250 1.6031 0.05572 0.04915 -0.0794 0.1635 1.0000
14.500 1.5995 0.05890 0.05232 -0.0787 0.1519 1.0000
14.750 1.5951 0.06227 0.05565 -0.0781 0.1415 1.0000
15.000 1.5908 0.06574 0.05912 -0.0777 0.1315 1.0000
15.250 1.5886 0.06905 0.06247 -0.0774 0.1219 1.0000
15.500 1.5850 0.07257 0.06601 -0.0771 0.1137 1.0000
15.750 1.5802 0.07632 0.06974 -0.0771 0.1063 1.0000
16.000 1.5788 0.07976 0.07327 -0.0771 0.0990 1.0000
16.250 1.5739 0.08360 0.07705 -0.0772 0.0931 1.0000
16.500 1.5731 0.08708 0.08066 -0.0775 0.0870 1.0000
16.750 1.5692 0.09086 0.08438 -0.0777 0.0821 1.0000
17.000 1.5686 0.09442 0.08809 -0.0781 0.0771 1.0000
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Polar data table (+)
Polar graphs
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