DAE-31 AIRFOIL (dae31-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: DAE-31 AIRFOIL (dae31-il) Reynolds number: 200,000 Max Cl/Cd: 91.28 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-dae31-il-200000.txt Download as CSV file: xf-dae31-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: DAE-31 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
------ -------- --------- --------- -------- -------- --------
-9.250 0.0565 0.09494 0.09102 -0.1137 0.9217 0.0371
-9.000 0.0594 0.09305 0.08913 -0.1156 0.9163 0.0379
-8.750 0.0598 0.09115 0.08723 -0.1181 0.9117 0.0383
-8.500 0.0755 0.08816 0.08425 -0.1169 0.9083 0.0389
-8.250 0.0834 0.08613 0.08223 -0.1171 0.9035 0.0395
-8.000 0.0910 0.08411 0.08022 -0.1175 0.8991 0.0404
-7.750 0.0965 0.08209 0.07818 -0.1183 0.8951 0.0418
-7.500 0.0864 0.08103 0.07719 -0.1205 0.8871 0.0429
-7.250 0.0908 0.07870 0.07488 -0.1207 0.8823 0.0432
-7.000 0.1084 0.07622 0.07238 -0.1194 0.8799 0.0443
-6.750 0.1134 0.07456 0.07073 -0.1192 0.8756 0.0458
-6.500 0.1095 0.07342 0.06965 -0.1191 0.8686 0.0473
-6.250 0.1076 0.07102 0.06725 -0.1264 0.8619 0.0484
-6.000 0.1211 0.06884 0.06506 -0.1228 0.8598 0.0491
-5.750 0.1274 0.06738 0.06364 -0.1224 0.8538 0.0501
-5.500 0.1367 0.06541 0.06167 -0.1237 0.8483 0.0516
-5.250 0.1570 0.06087 0.05702 -0.1359 0.8431 0.0544
-5.000 0.1688 0.05920 0.05536 -0.1328 0.8409 0.0552
-4.750 0.1742 0.05806 0.05427 -0.1318 0.8329 0.0562
-4.250 0.2222 0.05118 0.04719 -0.1416 0.8252 0.0621
-4.000 0.2447 0.04908 0.04504 -0.1426 0.8230 0.0648
-3.750 0.2632 0.04586 0.04165 -0.1472 0.8138 0.0694
-3.500 0.2972 0.04272 0.03811 -0.1521 0.8100 0.0773
-3.250 0.3178 0.04068 0.03620 -0.1514 0.8076 0.0789
-3.000 0.3503 0.03809 0.03334 -0.1542 0.8055 0.0882
-2.250 0.4387 0.02747 0.02085 -0.1574 0.7911 0.0542
-2.000 0.4696 0.02541 0.01870 -0.1578 0.7889 0.0523
-1.750 0.5026 0.02367 0.01657 -0.1581 0.7871 0.0509
-1.500 0.5204 0.02348 0.01621 -0.1567 0.7793 0.0528
-1.250 0.5488 0.02262 0.01510 -0.1563 0.7752 0.0546
-1.000 0.5796 0.02163 0.01409 -0.1564 0.7725 0.0568
-0.750 0.6117 0.02079 0.01314 -0.1565 0.7703 0.0603
-0.500 0.6438 0.02002 0.01238 -0.1567 0.7684 0.0675
-0.250 0.6601 0.02011 0.01259 -0.1551 0.7594 0.0734
0.000 0.6901 0.01946 0.01199 -0.1551 0.7559 0.0893
0.250 0.7222 0.01855 0.01133 -0.1555 0.7534 0.1531
0.500 0.7524 0.01727 0.01106 -0.1558 0.7514 0.4837
0.750 0.7745 0.01656 0.01133 -0.1546 0.7440 1.0000
1.000 0.8033 0.01645 0.01105 -0.1544 0.7394 1.0000
1.250 0.8354 0.01615 0.01059 -0.1546 0.7364 1.0000
1.500 0.8688 0.01581 0.01009 -0.1550 0.7340 1.0000
1.750 0.8888 0.01616 0.01041 -0.1538 0.7255 1.0000
2.000 0.9191 0.01592 0.01007 -0.1538 0.7212 1.0000
2.250 0.9522 0.01555 0.00958 -0.1542 0.7180 1.0000
2.500 0.9761 0.01567 0.00967 -0.1534 0.7109 1.0000
2.750 1.0045 0.01551 0.00945 -0.1532 0.7050 1.0000
3.000 1.0373 0.01514 0.00898 -0.1535 0.7011 1.0000
3.250 1.0611 0.01524 0.00907 -0.1527 0.6932 1.0000
3.500 1.0900 0.01506 0.00884 -0.1526 0.6870 1.0000
3.750 1.1233 0.01473 0.00840 -0.1531 0.6825 1.0000
4.000 1.1442 0.01492 0.00863 -0.1518 0.6726 1.0000
4.250 1.1764 0.01463 0.00826 -0.1522 0.6668 1.0000
4.500 1.1983 0.01477 0.00843 -0.1511 0.6566 1.0000
4.750 1.2297 0.01454 0.00811 -0.1513 0.6497 1.0000
5.000 1.2512 0.01470 0.00829 -0.1501 0.6385 1.0000
5.250 1.2781 0.01468 0.00824 -0.1497 0.6293 1.0000
5.500 1.3032 0.01471 0.00825 -0.1491 0.6186 1.0000
5.750 1.3260 0.01486 0.00840 -0.1481 0.6068 1.0000
6.000 1.3513 0.01494 0.00843 -0.1475 0.5958 1.0000
6.250 1.3756 0.01507 0.00853 -0.1468 0.5837 1.0000
6.500 1.3961 0.01532 0.00879 -0.1455 0.5703 1.0000
6.750 1.4174 0.01557 0.00902 -0.1443 0.5569 1.0000
7.000 1.4388 0.01584 0.00926 -0.1431 0.5434 1.0000
7.250 1.4594 0.01615 0.00951 -0.1419 0.5294 1.0000
7.500 1.4787 0.01651 0.00981 -0.1404 0.5150 1.0000
7.750 1.4947 0.01692 0.01022 -0.1385 0.4998 1.0000
8.000 1.5098 0.01738 0.01067 -0.1364 0.4847 1.0000
8.250 1.5236 0.01786 0.01114 -0.1341 0.4697 1.0000
8.500 1.5343 0.01839 0.01163 -0.1313 0.4551 1.0000
8.750 1.5440 0.01897 0.01219 -0.1284 0.4406 1.0000
9.000 1.5530 0.01964 0.01282 -0.1256 0.4260 1.0000
9.250 1.5614 0.02038 0.01351 -0.1228 0.4112 1.0000
9.500 1.5685 0.02120 0.01431 -0.1200 0.3962 1.0000
9.750 1.5753 0.02210 0.01520 -0.1172 0.3810 1.0000
10.000 1.5819 0.02308 0.01616 -0.1146 0.3660 1.0000
10.250 1.5882 0.02413 0.01721 -0.1121 0.3511 1.0000
10.500 1.5942 0.02526 0.01832 -0.1096 0.3365 1.0000
10.750 1.5997 0.02647 0.01951 -0.1072 0.3220 1.0000
11.000 1.6048 0.02777 0.02077 -0.1049 0.3079 1.0000
11.250 1.6092 0.02917 0.02213 -0.1026 0.2942 1.0000
11.500 1.6137 0.03061 0.02356 -0.1004 0.2806 1.0000
11.750 1.6182 0.03211 0.02507 -0.0984 0.2672 1.0000
12.000 1.6224 0.03368 0.02666 -0.0964 0.2544 1.0000
12.250 1.6260 0.03535 0.02832 -0.0944 0.2424 1.0000
12.500 1.6282 0.03718 0.03009 -0.0925 0.2311 1.0000
12.750 1.6320 0.03895 0.03191 -0.0909 0.2195 1.0000
13.000 1.6354 0.04083 0.03382 -0.0893 0.2085 1.0000
13.250 1.6376 0.04287 0.03585 -0.0877 0.1988 1.0000
13.500 1.6400 0.04497 0.03797 -0.0863 0.1889 1.0000
13.750 1.6434 0.04706 0.04013 -0.0851 0.1795 1.0000
14.000 1.6447 0.04938 0.04242 -0.0838 0.1713 1.0000
14.250 1.6475 0.05166 0.04479 -0.0828 0.1626 1.0000
14.500 1.6492 0.05410 0.04728 -0.0818 0.1547 1.0000
14.750 1.6497 0.05673 0.04991 -0.0809 0.1473 1.0000
15.000 1.6516 0.05930 0.05259 -0.0802 0.1396 1.0000
15.250 1.6501 0.06227 0.05552 -0.0794 0.1329 1.0000
15.500 1.6508 0.06516 0.05855 -0.0790 0.1254 1.0000
15.750 1.6489 0.06836 0.06179 -0.0786 0.1186 1.0000
16.000 1.6465 0.07175 0.06524 -0.0785 0.1119 1.0000
16.250 1.6446 0.07516 0.06875 -0.0784 0.1053 1.0000
16.500 1.6395 0.07902 0.07257 -0.0785 0.0995 1.0000
16.750 1.6374 0.08267 0.07639 -0.0788 0.0932 1.0000
17.000 1.6326 0.08668 0.08043 -0.0792 0.0878 1.0000
17.250 1.6279 0.09079 0.08463 -0.0798 0.0825 1.0000
17.500 1.6233 0.09495 0.08888 -0.0804 0.0774 1.0000
17.750 1.6170 0.09938 0.09331 -0.0813 0.0732 1.0000
18.000 1.6123 0.10369 0.09778 -0.0823 0.0685 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DAE-31 AIRFOIL (dae31-il)