DOUGLAS LA203A AIRFOIL (la203a-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: DOUGLAS LA203A AIRFOIL (la203a-il) Reynolds number: 100,000 Max Cl/Cd: 46.49 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-la203a-il-100000-n5.txt Download as CSV file: xf-la203a-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DOUGLAS LA203A AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.0242 0.09992 0.09377 -0.0971 0.8228 0.0778
-9.000 -0.0166 0.09724 0.09108 -0.0986 0.8175 0.0793
-8.750 -0.0122 0.09440 0.08822 -0.1009 0.8129 0.0813
-8.500 -0.0175 0.09105 0.08487 -0.1059 0.8079 0.0827
-8.250 -0.0205 0.08738 0.08122 -0.1107 0.8023 0.0831
-8.000 -0.0002 0.08498 0.07881 -0.1082 0.7985 0.0841
-7.750 0.0158 0.08294 0.07675 -0.1075 0.7946 0.0858
-7.500 0.0253 0.08047 0.07426 -0.1086 0.7907 0.0877
-7.000 0.0057 0.06478 0.05848 -0.1249 0.7802 0.0583
-6.750 0.0155 0.06129 0.05495 -0.1285 0.7764 0.0574
-6.500 0.0223 0.05673 0.05027 -0.1341 0.7724 0.0563
-6.250 0.0277 0.05187 0.04523 -0.1393 0.7673 0.0551
-6.000 0.0378 0.04692 0.03994 -0.1440 0.7629 0.0542
-5.750 0.0557 0.04335 0.03602 -0.1470 0.7595 0.0542
-5.500 0.0762 0.04089 0.03331 -0.1489 0.7559 0.0552
-5.250 0.0961 0.03855 0.03069 -0.1504 0.7512 0.0561
-5.000 0.1201 0.03615 0.02792 -0.1520 0.7473 0.0566
-4.750 0.1471 0.03393 0.02531 -0.1536 0.7441 0.0567
-4.500 0.1754 0.03203 0.02302 -0.1549 0.7410 0.0571
-4.250 0.1992 0.03067 0.02139 -0.1552 0.7360 0.0576
-4.000 0.2269 0.02935 0.01974 -0.1559 0.7322 0.0584
-3.750 0.2554 0.02840 0.01864 -0.1565 0.7291 0.0598
-3.500 0.2837 0.02772 0.01790 -0.1569 0.7261 0.0614
-3.250 0.3078 0.02715 0.01724 -0.1567 0.7212 0.0630
-3.000 0.3354 0.02639 0.01631 -0.1568 0.7175 0.0642
-2.750 0.3646 0.02564 0.01538 -0.1570 0.7146 0.0655
-2.500 0.3945 0.02493 0.01459 -0.1574 0.7122 0.0668
-2.250 0.4177 0.02469 0.01442 -0.1569 0.7074 0.0688
-2.000 0.4441 0.02438 0.01409 -0.1568 0.7034 0.0718
-1.750 0.4727 0.02396 0.01361 -0.1568 0.7003 0.0745
-1.500 0.5023 0.02347 0.01317 -0.1571 0.6978 0.0770
-1.250 0.5279 0.02329 0.01303 -0.1569 0.6938 0.0800
-1.000 0.5523 0.02324 0.01302 -0.1566 0.6889 0.0843
-0.750 0.5811 0.02302 0.01284 -0.1569 0.6854 0.0901
-0.500 0.6128 0.02272 0.01252 -0.1577 0.6826 0.0964
-0.250 0.6431 0.02255 0.01238 -0.1584 0.6793 0.1069
0.000 0.6670 0.02269 0.01266 -0.1584 0.6736 0.1236
0.250 0.7003 0.02218 0.01259 -0.1602 0.6698 0.2222
0.500 0.7308 0.02148 0.01323 -0.1607 0.6668 0.6501
0.750 0.7562 0.02159 0.01335 -0.1593 0.6644 0.7033
1.000 0.7689 0.02229 0.01419 -0.1565 0.6579 0.7391
1.250 0.7839 0.02260 0.01456 -0.1533 0.6536 0.7730
1.500 0.8045 0.02268 0.01461 -0.1511 0.6504 0.7967
1.750 0.8309 0.02264 0.01450 -0.1503 0.6479 0.8133
2.000 0.8427 0.02325 0.01519 -0.1477 0.6414 0.8244
2.250 0.8661 0.02349 0.01541 -0.1471 0.6369 0.8344
2.500 0.8937 0.02348 0.01535 -0.1469 0.6336 0.8418
2.750 0.9257 0.02340 0.01519 -0.1475 0.6311 0.8499
3.000 0.9391 0.02405 0.01593 -0.1455 0.6247 0.8587
3.250 0.9619 0.02437 0.01626 -0.1450 0.6197 0.8679
3.500 0.9877 0.02434 0.01621 -0.1444 0.6163 0.8771
3.750 1.0167 0.02419 0.01602 -0.1443 0.6138 0.8876
4.000 1.0265 0.02499 0.01694 -0.1419 0.6070 0.9014
4.250 1.0442 0.02526 0.01725 -0.1404 0.6019 0.9167
4.500 1.0683 0.02513 0.01712 -0.1394 0.5986 0.9367
4.750 1.0971 0.02483 0.01679 -0.1393 0.5961 0.9958
5.000 1.1118 0.02617 0.01824 -0.1389 0.5880 0.9999
5.250 1.1418 0.02647 0.01854 -0.1398 0.5835 0.9999
5.500 1.1776 0.02642 0.01846 -0.1412 0.5804 0.9999
5.750 1.2163 0.02624 0.01823 -0.1429 0.5781 0.9999
6.000 1.2160 0.02805 0.02020 -0.1402 0.5685 0.9999
6.250 1.2470 0.02813 0.02027 -0.1408 0.5646 0.9999
6.500 1.2844 0.02786 0.01999 -0.1421 0.5618 0.9999
6.750 1.2815 0.02961 0.02185 -0.1388 0.5529 0.9999
7.000 1.3072 0.02984 0.02211 -0.1387 0.5479 0.9999
7.250 1.3457 0.02939 0.02165 -0.1399 0.5447 0.9999
7.500 1.3391 0.03136 0.02374 -0.1363 0.5352 0.9999
7.750 1.3664 0.03136 0.02376 -0.1361 0.5299 0.9999
8.000 1.4110 0.03035 0.02274 -0.1377 0.5263 0.9999
8.250 1.3914 0.03311 0.02565 -0.1331 0.5147 0.9999
8.500 1.4346 0.03189 0.02440 -0.1340 0.5097 0.9999
8.750 1.4204 0.03454 0.02718 -0.1304 0.4981 0.9999
9.000 1.4651 0.03294 0.02553 -0.1311 0.4915 0.9999
9.250 1.4493 0.03589 0.02861 -0.1277 0.4791 0.9999
9.500 1.4628 0.03664 0.02939 -0.1264 0.4687 0.9999
9.750 1.4827 0.03681 0.02955 -0.1253 0.4578 0.9999
10.000 1.4755 0.03948 0.03232 -0.1230 0.4449 0.9999
10.250 1.4839 0.04078 0.03365 -0.1216 0.4326 0.9999
10.500 1.4994 0.04137 0.03424 -0.1203 0.4200 0.9999
10.750 1.5034 0.04311 0.03601 -0.1187 0.4063 0.9999
11.000 1.5016 0.04552 0.03847 -0.1170 0.3919 0.9999
11.250 1.5028 0.04764 0.04062 -0.1154 0.3770 0.9999
11.500 1.5047 0.04972 0.04271 -0.1139 0.3614 0.9999
11.750 1.5058 0.05188 0.04488 -0.1125 0.3446 0.9999
12.000 1.5059 0.05421 0.04718 -0.1110 0.3265 0.9999
12.250 1.5051 0.05669 0.04961 -0.1096 0.3071 0.9999
12.500 1.5035 0.05929 0.05212 -0.1083 0.2864 0.9999
12.750 1.4995 0.06228 0.05503 -0.1071 0.2652 0.9999
13.000 1.4942 0.06549 0.05814 -0.1059 0.2443 0.9999
13.250 1.4877 0.06895 0.06150 -0.1049 0.2247 0.9999
13.500 1.4806 0.07257 0.06505 -0.1041 0.2068 0.9999
13.750 1.4735 0.07630 0.06871 -0.1033 0.1903 0.9999
14.000 1.4667 0.08010 0.07246 -0.1028 0.1750 0.9999
14.250 1.4608 0.08389 0.07621 -0.1024 0.1612 0.9999
14.500 1.4556 0.08762 0.07992 -0.1021 0.1488 0.9999
14.750 1.4508 0.09136 0.08363 -0.1019 0.1378 0.9999
15.000 1.4460 0.09516 0.08740 -0.1019 0.1281 0.9999
15.250 1.4427 0.09877 0.09103 -0.1019 0.1190 0.9999
|
Polar data table (+)
Polar graphs
<< Back to DOUGLAS LA203A AIRFOIL (la203a-il)