DAE-21 AIRFOIL (dae21-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: DAE-21 AIRFOIL (dae21-il) Reynolds number: 1,000,000 Max Cl/Cd: 158.11 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-dae21-il-1000000-n5.txt Download as CSV file: xf-dae21-il-1000000-n5.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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.0770 0.09562 0.09262 -0.0712 0.7023 0.0081
-9.250 -0.0713 0.09235 0.08934 -0.0726 0.7001 0.0079
-9.000 -0.0664 0.08898 0.08596 -0.0741 0.6980 0.0078
-8.750 -0.0586 0.08640 0.08339 -0.0754 0.6962 0.0080
-8.500 -0.0518 0.08359 0.08059 -0.0767 0.6943 0.0081
-8.250 -0.0457 0.08070 0.07772 -0.0780 0.6920 0.0082
-8.000 -0.0399 0.07786 0.07489 -0.0794 0.6895 0.0083
-7.500 -0.1763 0.01960 0.01545 -0.1266 0.6843 0.0098
-7.250 -0.1505 0.01822 0.01387 -0.1273 0.6820 0.0099
-7.000 -0.1240 0.01712 0.01259 -0.1278 0.6799 0.0101
-6.750 -0.0968 0.01621 0.01154 -0.1282 0.6779 0.0102
-6.500 -0.0694 0.01538 0.01057 -0.1285 0.6756 0.0104
-6.250 -0.0419 0.01459 0.00963 -0.1288 0.6731 0.0105
-6.000 -0.0142 0.01383 0.00872 -0.1291 0.6706 0.0107
-5.750 0.0137 0.01313 0.00787 -0.1293 0.6680 0.0109
-5.500 0.0417 0.01253 0.00711 -0.1295 0.6655 0.0111
-5.250 0.0699 0.01197 0.00642 -0.1297 0.6632 0.0114
-5.000 0.0984 0.01148 0.00583 -0.1298 0.6608 0.0116
-4.750 0.1269 0.01108 0.00534 -0.1300 0.6582 0.0118
-4.500 0.1556 0.01080 0.00498 -0.1301 0.6555 0.0120
-4.250 0.1840 0.01035 0.00447 -0.1303 0.6527 0.0125
-4.000 0.2126 0.01012 0.00419 -0.1304 0.6499 0.0128
-3.750 0.2414 0.00988 0.00391 -0.1305 0.6474 0.0131
-3.500 0.2703 0.00963 0.00363 -0.1307 0.6447 0.0135
-3.250 0.2991 0.00940 0.00335 -0.1308 0.6417 0.0139
-3.000 0.3280 0.00919 0.00310 -0.1310 0.6386 0.0143
-2.750 0.3568 0.00902 0.00288 -0.1311 0.6356 0.0146
-2.500 0.3855 0.00882 0.00263 -0.1312 0.6325 0.0151
-2.250 0.4145 0.00864 0.00245 -0.1314 0.6298 0.0157
-2.000 0.4434 0.00850 0.00231 -0.1316 0.6265 0.0163
-1.750 0.4723 0.00840 0.00218 -0.1317 0.6230 0.0171
-1.500 0.5011 0.00832 0.00207 -0.1318 0.6196 0.0178
-1.250 0.5299 0.00820 0.00193 -0.1320 0.6163 0.0189
-1.000 0.5589 0.00810 0.00184 -0.1321 0.6130 0.0200
-0.750 0.5878 0.00803 0.00176 -0.1323 0.6090 0.0211
-0.500 0.6165 0.00796 0.00168 -0.1324 0.6051 0.0231
-0.250 0.6451 0.00792 0.00162 -0.1325 0.6013 0.0260
0.000 0.6741 0.00785 0.00158 -0.1327 0.5976 0.0307
0.250 0.7028 0.00780 0.00155 -0.1329 0.5933 0.0395
0.500 0.7314 0.00775 0.00154 -0.1330 0.5888 0.0574
0.750 0.7598 0.00768 0.00154 -0.1332 0.5846 0.0883
1.000 0.7885 0.00756 0.00156 -0.1334 0.5799 0.1420
1.250 0.8167 0.00745 0.00159 -0.1336 0.5745 0.2141
1.500 0.8447 0.00729 0.00166 -0.1338 0.5695 0.3230
1.750 0.8727 0.00709 0.00176 -0.1340 0.5640 0.4510
2.000 0.9004 0.00698 0.00184 -0.1341 0.5577 0.5458
2.250 0.9238 0.00645 0.00200 -0.1333 0.5522 0.8251
2.750 0.9832 0.00640 0.00210 -0.1341 0.5382 1.0000
3.000 1.0113 0.00650 0.00217 -0.1342 0.5305 1.0000
3.250 1.0388 0.00665 0.00225 -0.1342 0.5224 1.0000
3.500 1.0665 0.00677 0.00234 -0.1342 0.5138 1.0000
3.750 1.0937 0.00693 0.00245 -0.1342 0.5051 1.0000
4.000 1.1209 0.00709 0.00256 -0.1342 0.4957 1.0000
4.250 1.1479 0.00726 0.00269 -0.1342 0.4866 1.0000
4.500 1.1744 0.00746 0.00283 -0.1340 0.4764 1.0000
4.750 1.2011 0.00764 0.00298 -0.1340 0.4659 1.0000
5.000 1.2272 0.00786 0.00315 -0.1338 0.4546 1.0000
5.250 1.2528 0.00811 0.00334 -0.1336 0.4423 1.0000
5.500 1.2779 0.00838 0.00356 -0.1333 0.4290 1.0000
5.750 1.3028 0.00866 0.00377 -0.1330 0.4153 1.0000
6.000 1.3273 0.00895 0.00401 -0.1326 0.4016 1.0000
6.250 1.3511 0.00928 0.00427 -0.1322 0.3865 1.0000
6.500 1.3739 0.00965 0.00457 -0.1316 0.3702 1.0000
6.750 1.3957 0.01006 0.00490 -0.1308 0.3530 1.0000
7.000 1.4166 0.01050 0.00525 -0.1299 0.3355 1.0000
7.250 1.4365 0.01097 0.00564 -0.1289 0.3172 1.0000
7.500 1.4555 0.01144 0.00604 -0.1277 0.3004 1.0000
7.750 1.4727 0.01197 0.00648 -0.1263 0.2830 1.0000
8.000 1.4887 0.01250 0.00694 -0.1246 0.2679 1.0000
8.250 1.4996 0.01308 0.00746 -0.1220 0.2525 1.0000
8.500 1.5067 0.01374 0.00808 -0.1190 0.2392 1.0000
8.750 1.5133 0.01457 0.00884 -0.1162 0.2247 1.0000
9.000 1.5207 0.01550 0.00971 -0.1137 0.2105 1.0000
9.250 1.5289 0.01648 0.01064 -0.1115 0.1971 1.0000
9.500 1.5374 0.01751 0.01162 -0.1095 0.1844 1.0000
9.750 1.5462 0.01858 0.01264 -0.1076 0.1721 1.0000
10.000 1.5538 0.01976 0.01377 -0.1057 0.1589 1.0000
10.250 1.5620 0.02095 0.01492 -0.1040 0.1472 1.0000
10.500 1.5699 0.02219 0.01613 -0.1022 0.1361 1.0000
10.750 1.5773 0.02349 0.01739 -0.1005 0.1255 1.0000
11.000 1.5843 0.02486 0.01872 -0.0989 0.1160 1.0000
11.250 1.5909 0.02629 0.02012 -0.0973 0.1061 1.0000
11.500 1.5984 0.02768 0.02149 -0.0958 0.0976 1.0000
11.750 1.6054 0.02914 0.02294 -0.0943 0.0902 1.0000
12.000 1.6111 0.03073 0.02449 -0.0929 0.0822 1.0000
12.250 1.6182 0.03223 0.02601 -0.0916 0.0761 1.0000
12.500 1.6254 0.03379 0.02756 -0.0903 0.0706 1.0000
12.750 1.6311 0.03552 0.02929 -0.0891 0.0647 1.0000
13.000 1.6390 0.03710 0.03088 -0.0881 0.0604 1.0000
13.250 1.6439 0.03901 0.03278 -0.0870 0.0551 1.0000
13.500 1.6520 0.04065 0.03446 -0.0861 0.0517 1.0000
13.750 1.6577 0.04256 0.03638 -0.0852 0.0480 1.0000
14.000 1.6644 0.04442 0.03826 -0.0843 0.0447 1.0000
14.250 1.6701 0.04641 0.04028 -0.0836 0.0417 1.0000
14.500 1.6755 0.04848 0.04238 -0.0828 0.0386 1.0000
14.750 1.6814 0.05054 0.04447 -0.0822 0.0363 1.0000
15.000 1.6856 0.05281 0.04677 -0.0816 0.0339 1.0000
15.250 1.6913 0.05497 0.04897 -0.0811 0.0319 1.0000
15.500 1.6950 0.05739 0.05143 -0.0807 0.0299 1.0000
15.750 1.6997 0.05973 0.05381 -0.0803 0.0283 1.0000
16.000 1.7037 0.06220 0.05635 -0.0800 0.0269 1.0000
16.250 1.7068 0.06482 0.05901 -0.0797 0.0253 1.0000
16.500 1.7094 0.06755 0.06179 -0.0796 0.0239 1.0000
16.750 1.7130 0.07019 0.06449 -0.0795 0.0228 1.0000
17.000 1.7139 0.07321 0.06757 -0.0795 0.0215 1.0000
17.250 1.7154 0.07623 0.07064 -0.0796 0.0206 1.0000
17.500 1.7179 0.07913 0.07362 -0.0798 0.0198 1.0000
17.750 1.7187 0.08234 0.07690 -0.0801 0.0190 1.0000
18.000 1.7191 0.08562 0.08025 -0.0804 0.0181 1.0000
18.250 1.7177 0.08921 0.08390 -0.0809 0.0173 1.0000
18.500 1.7179 0.09262 0.08739 -0.0815 0.0168 1.0000
18.750 1.7180 0.09605 0.09090 -0.0821 0.0161 1.0000
19.000 1.7163 0.09985 0.09477 -0.0829 0.0155 1.0000
19.250 1.7138 0.10377 0.09877 -0.0838 0.0150 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DAE-21 AIRFOIL (dae21-il)