EPPLER 561 AIRFOIL (e561-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: EPPLER 561 AIRFOIL (e561-il) Reynolds number: 100,000 Max Cl/Cd: 39.8 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e561-il-100000.txt Download as CSV file: xf-e561-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 561 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6051 0.06415 0.05939 -0.0708 1.0000 0.0614
-10.000 -0.6382 0.06229 0.05760 -0.0680 1.0000 0.0611
-9.750 -0.6716 0.05888 0.05415 -0.0688 1.0000 0.0606
-9.500 -0.6809 0.05248 0.04730 -0.0773 0.9945 0.0606
-9.250 -0.6581 0.04580 0.03976 -0.0883 0.9840 0.0616
-9.000 -0.6272 0.04247 0.03631 -0.0923 0.9756 0.0632
-8.750 -0.5923 0.03986 0.03351 -0.0964 0.9673 0.0654
-8.500 -0.5531 0.03712 0.03044 -0.1010 0.9601 0.0684
-8.250 -0.5205 0.03471 0.02780 -0.1037 0.9505 0.0720
-8.000 -0.4836 0.03310 0.02615 -0.1067 0.9419 0.0784
-7.750 -0.4425 0.03126 0.02430 -0.1106 0.9344 0.0890
-7.500 -0.4088 0.02943 0.02252 -0.1131 0.9244 0.1058
-7.250 -0.3600 0.02804 0.02115 -0.1185 0.9183 0.1429
-7.000 -0.3268 0.02786 0.02111 -0.1201 0.9069 0.1668
-6.750 -0.2759 0.02800 0.02123 -0.1243 0.9018 0.1937
-6.500 -0.2452 0.02871 0.02206 -0.1244 0.8896 0.2124
-6.250 -0.1947 0.02885 0.02221 -0.1277 0.8855 0.2325
-6.000 -0.1650 0.02868 0.02201 -0.1278 0.8737 0.2451
-5.750 -0.1137 0.02819 0.02144 -0.1315 0.8698 0.2612
-5.500 -0.0802 0.02771 0.02079 -0.1326 0.8590 0.2746
-5.250 -0.0293 0.02704 0.01995 -0.1365 0.8539 0.2902
-5.000 0.0061 0.02704 0.02005 -0.1370 0.8444 0.2999
-4.750 0.0511 0.02634 0.01920 -0.1399 0.8371 0.3127
-4.500 0.0882 0.02568 0.01830 -0.1418 0.8271 0.3257
-4.250 0.1294 0.02537 0.01800 -0.1434 0.8187 0.3360
-4.000 0.1605 0.02493 0.01741 -0.1441 0.8072 0.3466
-3.750 0.2046 0.02429 0.01647 -0.1471 0.7992 0.3604
-3.500 0.2286 0.02426 0.01653 -0.1459 0.7865 0.3676
-3.250 0.2636 0.02380 0.01583 -0.1474 0.7767 0.3799
-3.000 0.2931 0.02368 0.01571 -0.1472 0.7661 0.3886
-2.750 0.3213 0.02342 0.01528 -0.1475 0.7551 0.3994
-2.500 0.3556 0.02320 0.01496 -0.1483 0.7463 0.4097
-2.250 0.3792 0.02311 0.01479 -0.1476 0.7347 0.4194
-2.000 0.4143 0.02290 0.01440 -0.1488 0.7269 0.4311
-1.750 0.4360 0.02291 0.01443 -0.1476 0.7157 0.4397
-1.500 0.4687 0.02277 0.01409 -0.1485 0.7076 0.4524
-1.250 0.4918 0.02279 0.01417 -0.1475 0.6974 0.4608
-1.000 0.5230 0.02270 0.01393 -0.1481 0.6898 0.4728
-0.750 0.5464 0.02280 0.01405 -0.1473 0.6804 0.4832
-0.500 0.5792 0.02270 0.01383 -0.1480 0.6738 0.4947
-0.250 0.6009 0.02289 0.01400 -0.1471 0.6640 0.5068
0.000 0.6322 0.02282 0.01389 -0.1474 0.6576 0.5183
0.250 0.6531 0.02307 0.01419 -0.1464 0.6489 0.5294
0.500 0.6831 0.02310 0.01413 -0.1467 0.6421 0.5434
0.750 0.7077 0.02328 0.01436 -0.1461 0.6351 0.5554
1.000 0.7318 0.02344 0.01456 -0.1454 0.6274 0.5681
1.250 0.7641 0.02345 0.01448 -0.1460 0.6221 0.5834
1.500 0.7828 0.02388 0.01502 -0.1447 0.6143 0.5978
1.750 0.8100 0.02401 0.01514 -0.1445 0.6080 0.6141
2.000 0.8401 0.02411 0.01522 -0.1447 0.6028 0.6318
2.250 0.8563 0.02459 0.01588 -0.1429 0.5952 0.6481
2.500 0.8841 0.02472 0.01602 -0.1428 0.5898 0.6691
2.750 0.9106 0.02490 0.01625 -0.1423 0.5848 0.6914
3.000 0.9253 0.02542 0.01698 -0.1402 0.5775 0.7154
3.250 0.9507 0.02551 0.01713 -0.1395 0.5723 0.7471
3.500 0.9764 0.02557 0.01725 -0.1386 0.5682 0.7890
3.750 0.9776 0.02615 0.01818 -0.1339 0.5614 0.8434
4.000 1.0033 0.02607 0.01818 -0.1335 0.5559 1.0000
4.250 1.0468 0.02630 0.01814 -0.1367 0.5516 1.0000
4.500 1.0574 0.02754 0.01950 -0.1349 0.5445 1.0000
4.750 1.0844 0.02806 0.01996 -0.1350 0.5390 1.0000
5.000 1.1216 0.02821 0.01995 -0.1365 0.5348 1.0000
5.250 1.1294 0.02957 0.02144 -0.1341 0.5282 1.0000
5.500 1.1514 0.03028 0.02217 -0.1335 0.5226 1.0000
5.750 1.1865 0.03044 0.02222 -0.1345 0.5184 1.0000
6.000 1.1975 0.03171 0.02359 -0.1325 0.5125 1.0000
6.250 1.2107 0.03282 0.02479 -0.1307 0.5064 1.0000
6.500 1.2441 0.03300 0.02490 -0.1315 0.5021 1.0000
6.750 1.2679 0.03374 0.02565 -0.1311 0.4973 1.0000
7.000 1.2607 0.03578 0.02790 -0.1268 0.4901 1.0000
7.250 1.2938 0.03594 0.02803 -0.1275 0.4857 1.0000
7.500 1.3406 0.03562 0.02758 -0.1299 0.4823 1.0000
7.750 1.2965 0.03943 0.03174 -0.1213 0.4736 1.0000
8.000 1.3316 0.03943 0.03173 -0.1222 0.4692 1.0000
8.250 1.3838 0.03871 0.03092 -0.1250 0.4658 1.0000
8.500 1.2593 0.04687 0.03947 -0.1086 0.4547 1.0000
8.750 1.3360 0.04437 0.03693 -0.1132 0.4520 1.0000
9.000 1.4156 0.04228 0.03475 -0.1189 0.4491 1.0000
9.250 1.1970 0.05916 0.05197 -0.0998 0.4326 1.0000
9.500 1.2945 0.05296 0.04576 -0.1022 0.4331 1.0000
9.750 1.4170 0.04708 0.03982 -0.1096 0.4321 1.0000
10.000 0.9511 0.10044 0.09346 -0.1027 0.3858 1.0000
10.250 0.9888 0.09927 0.09233 -0.1017 0.3828 1.0000
10.500 1.0383 0.09646 0.08954 -0.1002 0.3812 1.0000
10.750 0.9753 0.10853 0.10169 -0.1030 0.3677 1.0000
11.000 1.0199 0.10629 0.09948 -0.1015 0.3653 1.0000
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Polar data table (+)
Polar graphs
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