EPPLER 654 AIRFOIL (e654-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 654 AIRFOIL (e654-il) Reynolds number: 1,000,000 Max Cl/Cd: 158.57 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e654-il-1000000.txt Download as CSV file: xf-e654-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 654 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.500 -0.2851 0.06722 0.06395 -0.1307 0.7495 0.0082
-13.250 -0.3008 0.06015 0.05676 -0.1356 0.7462 0.0081
-13.000 -0.3217 0.05331 0.04981 -0.1399 0.7427 0.0079
-12.750 -0.3285 0.04901 0.04540 -0.1429 0.7385 0.0079
-12.500 -0.3657 0.04166 0.03784 -0.1470 0.7352 0.0078
-12.250 -0.3788 0.03802 0.03406 -0.1484 0.7320 0.0078
-12.000 -0.3912 0.03486 0.03076 -0.1487 0.7288 0.0077
-11.750 -0.4040 0.03197 0.02770 -0.1482 0.7254 0.0076
-11.500 -0.4140 0.02950 0.02504 -0.1470 0.7222 0.0075
-11.250 -0.4244 0.02706 0.02239 -0.1451 0.7188 0.0075
-11.000 -0.4263 0.02534 0.02053 -0.1432 0.7167 0.0074
-10.750 -0.4252 0.02402 0.01909 -0.1410 0.7142 0.0073
-10.500 -0.4281 0.02263 0.01754 -0.1380 0.7117 0.0073
-10.250 -0.4193 0.02136 0.01612 -0.1363 0.7091 0.0072
-10.000 -0.4067 0.02028 0.01488 -0.1349 0.7066 0.0072
-9.750 -0.3917 0.01928 0.01374 -0.1338 0.7039 0.0072
-9.500 -0.3744 0.01841 0.01275 -0.1329 0.7016 0.0072
-9.250 -0.3561 0.01751 0.01176 -0.1320 0.7001 0.0072
-9.000 -0.3360 0.01675 0.01092 -0.1314 0.6984 0.0072
-8.750 -0.3151 0.01604 0.01013 -0.1308 0.6966 0.0072
-8.500 -0.2934 0.01538 0.00940 -0.1303 0.6947 0.0073
-8.250 -0.2707 0.01480 0.00874 -0.1299 0.6927 0.0074
-8.000 -0.2473 0.01426 0.00813 -0.1296 0.6907 0.0075
-7.750 -0.2232 0.01378 0.00758 -0.1293 0.6888 0.0076
-7.500 -0.1984 0.01335 0.00708 -0.1292 0.6866 0.0077
-7.250 -0.1729 0.01296 0.00663 -0.1291 0.6847 0.0078
-7.000 -0.1468 0.01258 0.00622 -0.1291 0.6835 0.0080
-6.750 -0.1217 0.01201 0.00560 -0.1291 0.6821 0.0085
-6.500 -0.0950 0.01165 0.00522 -0.1292 0.6804 0.0091
-6.250 -0.0677 0.01137 0.00492 -0.1293 0.6787 0.0099
-6.000 -0.0405 0.01102 0.00455 -0.1295 0.6770 0.0114
-5.750 -0.0130 0.01072 0.00424 -0.1297 0.6754 0.0144
-5.500 0.0148 0.01045 0.00398 -0.1299 0.6739 0.0198
-5.250 0.0428 0.01021 0.00376 -0.1302 0.6722 0.0277
-5.000 0.0711 0.01003 0.00360 -0.1306 0.6703 0.0361
-4.750 0.0997 0.00988 0.00347 -0.1310 0.6681 0.0468
-4.500 0.1281 0.00965 0.00331 -0.1314 0.6670 0.0606
-4.250 0.1567 0.00942 0.00317 -0.1318 0.6657 0.0786
-4.000 0.1855 0.00919 0.00303 -0.1323 0.6643 0.1027
-3.750 0.2143 0.00891 0.00290 -0.1329 0.6628 0.1415
-3.500 0.2431 0.00856 0.00277 -0.1336 0.6612 0.2006
-3.250 0.2723 0.00820 0.00264 -0.1343 0.6595 0.2706
-3.000 0.3018 0.00789 0.00254 -0.1351 0.6579 0.3375
-2.750 0.3316 0.00762 0.00245 -0.1359 0.6564 0.4041
-2.500 0.3617 0.00735 0.00238 -0.1367 0.6547 0.4777
-2.250 0.3919 0.00715 0.00238 -0.1376 0.6527 0.5565
-2.000 0.4219 0.00710 0.00244 -0.1382 0.6510 0.6019
-1.750 0.4515 0.00707 0.00246 -0.1386 0.6497 0.6285
-1.500 0.4811 0.00707 0.00248 -0.1391 0.6483 0.6456
-1.250 0.5106 0.00709 0.00251 -0.1395 0.6466 0.6596
-1.000 0.5402 0.00711 0.00253 -0.1399 0.6448 0.6700
-0.750 0.5697 0.00714 0.00255 -0.1403 0.6429 0.6786
-0.500 0.5992 0.00718 0.00255 -0.1407 0.6410 0.6859
-0.250 0.6284 0.00721 0.00259 -0.1411 0.6391 0.6931
0.000 0.6580 0.00731 0.00264 -0.1415 0.6367 0.6999
0.250 0.6873 0.00738 0.00270 -0.1419 0.6344 0.7057
0.500 0.7161 0.00739 0.00274 -0.1422 0.6326 0.7111
0.750 0.7452 0.00741 0.00277 -0.1426 0.6304 0.7165
1.000 0.7743 0.00744 0.00279 -0.1429 0.6281 0.7212
1.250 0.8030 0.00745 0.00281 -0.1432 0.6257 0.7258
1.500 0.8318 0.00748 0.00285 -0.1435 0.6232 0.7301
1.750 0.8608 0.00757 0.00289 -0.1439 0.6204 0.7343
2.000 0.8898 0.00764 0.00296 -0.1443 0.6178 0.7378
2.250 0.9180 0.00762 0.00299 -0.1445 0.6153 0.7418
2.500 0.9464 0.00764 0.00303 -0.1448 0.6124 0.7453
2.750 0.9748 0.00766 0.00307 -0.1450 0.6095 0.7487
3.000 1.0032 0.00771 0.00310 -0.1453 0.6065 0.7519
3.250 1.0317 0.00781 0.00316 -0.1456 0.6029 0.7547
3.500 1.0596 0.00780 0.00321 -0.1458 0.6000 0.7576
3.750 1.0875 0.00780 0.00326 -0.1460 0.5964 0.7604
4.000 1.1151 0.00784 0.00331 -0.1462 0.5925 0.7632
4.250 1.1424 0.00791 0.00337 -0.1463 0.5883 0.7661
4.500 1.1700 0.00796 0.00345 -0.1464 0.5841 0.7689
4.750 1.1975 0.00800 0.00351 -0.1466 0.5791 0.7716
5.000 1.2240 0.00807 0.00357 -0.1465 0.5737 0.7743
5.250 1.2508 0.00812 0.00367 -0.1465 0.5684 0.7771
5.500 1.2772 0.00818 0.00376 -0.1465 0.5619 0.7798
5.750 1.3023 0.00830 0.00387 -0.1462 0.5553 0.7828
6.000 1.3284 0.00838 0.00398 -0.1461 0.5474 0.7858
6.250 1.3526 0.00853 0.00411 -0.1457 0.5388 0.7886
6.500 1.3763 0.00868 0.00426 -0.1451 0.5280 0.7914
6.750 1.3987 0.00885 0.00443 -0.1444 0.5155 0.7946
7.000 1.4193 0.00908 0.00465 -0.1433 0.5016 0.7979
7.250 1.4376 0.00938 0.00490 -0.1417 0.4851 0.8016
7.500 1.4526 0.00975 0.00521 -0.1396 0.4665 0.8052
8.000 1.4692 0.01064 0.00599 -0.1328 0.4285 0.8132
8.250 1.4768 0.01119 0.00648 -0.1294 0.4096 0.8177
8.500 1.4840 0.01180 0.00703 -0.1261 0.3922 0.8224
8.750 1.4889 0.01249 0.00768 -0.1226 0.3744 0.8276
9.000 1.4929 0.01328 0.00844 -0.1191 0.3577 0.8334
9.250 1.4965 0.01419 0.00930 -0.1158 0.3412 0.8393
9.500 1.4989 0.01521 0.01030 -0.1125 0.3252 0.8457
9.750 1.5023 0.01631 0.01138 -0.1096 0.3105 0.8530
10.000 1.5048 0.01754 0.01259 -0.1067 0.2955 0.8608
10.250 1.5077 0.01883 0.01387 -0.1041 0.2817 0.8705
10.500 1.5096 0.02021 0.01525 -0.1015 0.2679 0.8828
10.750 1.5105 0.02166 0.01671 -0.0988 0.2539 0.9007
11.000 1.5111 0.02296 0.01811 -0.0960 0.2413 1.0000
11.250 1.5167 0.02446 0.01956 -0.0944 0.2284 1.0000
11.500 1.5221 0.02600 0.02106 -0.0928 0.2161 1.0000
11.750 1.5263 0.02766 0.02268 -0.0911 0.2039 1.0000
12.000 1.5301 0.02937 0.02436 -0.0895 0.1922 1.0000
12.250 1.5343 0.03111 0.02605 -0.0880 0.1805 1.0000
12.500 1.5401 0.03278 0.02769 -0.0867 0.1699 1.0000
12.750 1.5463 0.03443 0.02933 -0.0855 0.1600 1.0000
13.000 1.5504 0.03633 0.03118 -0.0843 0.1502 1.0000
13.250 1.5566 0.03807 0.03291 -0.0832 0.1406 1.0000
13.500 1.5628 0.03987 0.03470 -0.0822 0.1323 1.0000
13.750 1.5679 0.04178 0.03657 -0.0812 0.1235 1.0000
14.000 1.5733 0.04372 0.03850 -0.0804 0.1153 1.0000
14.250 1.5794 0.04564 0.04040 -0.0796 0.1073 1.0000
14.750 1.5897 0.04976 0.04451 -0.0781 0.0929 1.0000
15.000 1.5941 0.05194 0.04667 -0.0775 0.0856 1.0000
15.250 1.5982 0.05422 0.04895 -0.0769 0.0796 1.0000
15.500 1.6039 0.05636 0.05110 -0.0764 0.0735 1.0000
15.750 1.6056 0.05899 0.05370 -0.0759 0.0672 1.0000
16.000 1.6114 0.06118 0.05592 -0.0756 0.0625 1.0000
16.250 1.6139 0.06382 0.05856 -0.0753 0.0578 1.0000
16.500 1.6178 0.06632 0.06108 -0.0751 0.0531 1.0000
16.750 1.6204 0.06904 0.06381 -0.0749 0.0489 1.0000
17.000 1.6229 0.07180 0.06658 -0.0748 0.0450 1.0000
17.250 1.6252 0.07462 0.06944 -0.0748 0.0414 1.0000
17.500 1.6267 0.07760 0.07243 -0.0749 0.0382 1.0000
17.750 1.6291 0.08053 0.07541 -0.0751 0.0357 1.0000
18.000 1.6300 0.08366 0.07857 -0.0753 0.0328 1.0000
18.250 1.6310 0.08685 0.08181 -0.0757 0.0307 1.0000
18.500 1.6325 0.08998 0.08498 -0.0761 0.0286 1.0000
18.750 1.6311 0.09358 0.08863 -0.0767 0.0266 1.0000
19.000 1.6323 0.09684 0.09195 -0.0773 0.0249 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 654 AIRFOIL (e654-il)