EPPLER 545 AIRFOIL (e545-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 545 AIRFOIL (e545-il) Reynolds number: 100,000 Max Cl/Cd: 42.86 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e545-il-100000-n5.txt Download as CSV file: xf-e545-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 545 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
------ -------- --------- --------- -------- -------- --------
-13.250 -0.4797 0.09698 0.09169 -0.0621 1.0000 0.0154
-13.000 -0.5035 0.08829 0.08292 -0.0669 1.0000 0.0152
-12.750 -0.5273 0.08127 0.07580 -0.0705 1.0000 0.0150
-12.500 -0.5514 0.07525 0.06964 -0.0728 1.0000 0.0148
-12.250 -0.5738 0.07015 0.06440 -0.0740 1.0000 0.0147
-12.000 -0.5953 0.06567 0.05975 -0.0743 1.0000 0.0146
-11.750 -0.6161 0.06165 0.05555 -0.0737 1.0000 0.0146
-11.500 -0.6326 0.05836 0.05209 -0.0725 1.0000 0.0145
-11.250 -0.6491 0.05532 0.04887 -0.0706 1.0000 0.0145
-11.000 -0.6634 0.05275 0.04614 -0.0682 1.0000 0.0145
-10.750 -0.6775 0.05052 0.04374 -0.0651 1.0000 0.0146
-10.500 -0.6932 0.04865 0.04172 -0.0612 1.0000 0.0147
-10.250 -0.7098 0.04741 0.04038 -0.0566 1.0000 0.0146
-10.000 -0.6975 0.04426 0.03682 -0.0576 0.9929 0.0152
-9.750 -0.6766 0.04124 0.03331 -0.0588 0.9843 0.0159
-9.500 -0.6545 0.03969 0.03171 -0.0605 0.9744 0.0168
-9.250 -0.6258 0.03750 0.02916 -0.0620 0.9676 0.0184
-9.000 -0.5979 0.03574 0.02728 -0.0631 0.9593 0.0196
-8.750 -0.5617 0.03403 0.02542 -0.0649 0.9549 0.0205
-8.500 -0.5298 0.03251 0.02371 -0.0656 0.9478 0.0218
-8.250 -0.4995 0.03114 0.02234 -0.0671 0.9414 0.0232
-8.000 -0.4717 0.02986 0.02099 -0.0683 0.9333 0.0264
-7.750 -0.4430 0.02863 0.01970 -0.0698 0.9256 0.0305
-7.500 -0.4189 0.02740 0.01842 -0.0705 0.9162 0.0346
-7.250 -0.3936 0.02625 0.01722 -0.0714 0.9074 0.0425
-7.000 -0.3751 0.02517 0.01618 -0.0711 0.8962 0.0541
-6.750 -0.3540 0.02397 0.01512 -0.0713 0.8875 0.0770
-6.500 -0.3450 0.02297 0.01434 -0.0693 0.8747 0.1120
-6.250 -0.3383 0.02168 0.01341 -0.0675 0.8639 0.1761
-6.000 -0.3411 0.02002 0.01229 -0.0645 0.8523 0.2765
-5.750 -0.3285 0.02075 0.01451 -0.0592 0.8432 0.5041
-5.500 -0.3104 0.02067 0.01419 -0.0582 0.8343 0.5618
-5.250 -0.2931 0.02082 0.01410 -0.0567 0.8252 0.5914
-5.000 -0.2633 0.02176 0.01485 -0.0558 0.8180 0.6107
-4.750 -0.2332 0.02246 0.01534 -0.0552 0.8119 0.6272
-4.500 -0.2093 0.02343 0.01618 -0.0532 0.8032 0.6440
-4.250 -0.1633 0.02526 0.01792 -0.0527 0.7982 0.6535
-4.000 -0.1449 0.02499 0.01747 -0.0517 0.7912 0.6639
-3.750 -0.1181 0.02505 0.01739 -0.0512 0.7838 0.6666
-3.500 -0.0892 0.02492 0.01709 -0.0514 0.7779 0.6699
-3.250 -0.0692 0.02464 0.01667 -0.0506 0.7705 0.6747
-3.000 -0.0533 0.02398 0.01584 -0.0500 0.7631 0.6813
-2.750 -0.0236 0.02384 0.01556 -0.0503 0.7582 0.6832
-2.500 -0.0013 0.02377 0.01542 -0.0495 0.7512 0.6854
-2.250 0.0231 0.02360 0.01513 -0.0492 0.7449 0.6879
-2.000 0.0503 0.02333 0.01473 -0.0495 0.7400 0.6908
-1.750 0.0698 0.02306 0.01437 -0.0489 0.7331 0.6945
-1.500 0.0915 0.02262 0.01380 -0.0491 0.7270 0.6985
-1.250 0.1193 0.02244 0.01352 -0.0494 0.7224 0.7004
-1.000 0.1432 0.02236 0.01340 -0.0489 0.7167 0.7021
-0.750 0.1664 0.02226 0.01325 -0.0485 0.7104 0.7038
-0.500 0.1934 0.02212 0.01303 -0.0487 0.7056 0.7056
-0.250 0.2216 0.02199 0.01282 -0.0492 0.7016 0.7078
0.000 0.2419 0.02194 0.01277 -0.0485 0.6951 0.7104
0.250 0.2674 0.02179 0.01256 -0.0487 0.6898 0.7127
0.500 0.2966 0.02159 0.01227 -0.0496 0.6855 0.7149
0.750 0.3208 0.02148 0.01212 -0.0498 0.6799 0.7174
1.000 0.3439 0.02148 0.01215 -0.0492 0.6737 0.7190
1.250 0.3717 0.02141 0.01205 -0.0493 0.6687 0.7206
1.500 0.3980 0.02138 0.01200 -0.0494 0.6636 0.7223
1.750 0.4196 0.02141 0.01208 -0.0487 0.6568 0.7241
2.000 0.4471 0.02135 0.01200 -0.0489 0.6515 0.7259
2.250 0.4770 0.02126 0.01187 -0.0497 0.6471 0.7277
2.500 0.4970 0.02136 0.01204 -0.0489 0.6403 0.7299
2.750 0.5236 0.02134 0.01203 -0.0491 0.6349 0.7323
3.000 0.5552 0.02124 0.01189 -0.0502 0.6303 0.7349
3.250 0.5748 0.02138 0.01212 -0.0492 0.6232 0.7369
3.500 0.5994 0.02140 0.01220 -0.0488 0.6171 0.7386
3.750 0.6296 0.02133 0.01214 -0.0494 0.6123 0.7403
4.000 0.6483 0.02150 0.01242 -0.0482 0.6048 0.7423
4.250 0.6740 0.02150 0.01248 -0.0480 0.5983 0.7444
4.500 0.7027 0.02145 0.01246 -0.0484 0.5924 0.7466
4.750 0.7219 0.02159 0.01271 -0.0474 0.5839 0.7491
5.000 0.7517 0.02151 0.01264 -0.0480 0.5774 0.7517
5.250 0.7717 0.02169 0.01294 -0.0471 0.5697 0.7544
5.500 0.7947 0.02177 0.01313 -0.0464 0.5630 0.7566
5.750 0.8221 0.02178 0.01321 -0.0464 0.5574 0.7589
6.000 0.8380 0.02202 0.01363 -0.0447 0.5486 0.7617
6.250 0.8666 0.02194 0.01358 -0.0449 0.5410 0.7643
6.500 0.8817 0.02218 0.01398 -0.0432 0.5307 0.7673
6.750 0.9076 0.02216 0.01401 -0.0430 0.5216 0.7703
7.000 0.9232 0.02236 0.01437 -0.0412 0.5106 0.7730
7.250 0.9411 0.02248 0.01462 -0.0397 0.5002 0.7757
7.500 0.9608 0.02255 0.01479 -0.0384 0.4893 0.7789
7.750 0.9727 0.02283 0.01522 -0.0361 0.4765 0.7828
8.000 0.9858 0.02304 0.01552 -0.0339 0.4629 0.7871
8.250 0.9978 0.02328 0.01585 -0.0315 0.4483 0.7910
8.500 1.0080 0.02360 0.01626 -0.0289 0.4320 0.7949
8.750 1.0177 0.02402 0.01674 -0.0263 0.4135 0.7992
9.000 1.0272 0.02454 0.01727 -0.0240 0.3931 0.8040
9.250 1.0356 0.02515 0.01787 -0.0216 0.3719 0.8087
9.500 1.0411 0.02592 0.01863 -0.0190 0.3506 0.8135
9.750 1.0452 0.02688 0.01956 -0.0165 0.3282 0.8193
10.000 1.0494 0.02797 0.02061 -0.0142 0.3080 0.8256
10.250 1.0523 0.02912 0.02177 -0.0119 0.2899 0.8323
10.500 1.0556 0.03042 0.02307 -0.0099 0.2725 0.8404
10.750 1.0575 0.03178 0.02448 -0.0078 0.2557 0.8490
11.000 1.0594 0.03327 0.02600 -0.0060 0.2392 0.8595
11.250 1.0600 0.03483 0.02760 -0.0042 0.2236 0.8726
11.500 1.0603 0.03649 0.02930 -0.0025 0.2089 0.8909
11.750 1.0622 0.03821 0.03112 -0.0015 0.1939 0.9290
12.000 1.0635 0.04003 0.03301 -0.0008 0.1779 1.0000
12.250 1.0687 0.04210 0.03517 -0.0007 0.1545 1.0000
12.500 1.0667 0.04475 0.03769 -0.0004 0.1368 1.0000
12.750 1.0631 0.04761 0.04039 -0.0001 0.1257 1.0000
13.000 1.0621 0.05034 0.04306 0.0001 0.1155 1.0000
13.250 1.0636 0.05291 0.04568 0.0001 0.1058 1.0000
13.500 1.0654 0.05552 0.04835 0.0001 0.0968 1.0000
13.750 1.0667 0.05827 0.05114 0.0000 0.0888 1.0000
14.000 1.0667 0.06123 0.05412 -0.0003 0.0820 1.0000
14.250 1.0679 0.06414 0.05712 -0.0007 0.0752 1.0000
14.500 1.0682 0.06723 0.06026 -0.0012 0.0695 1.0000
14.750 1.0664 0.07063 0.06369 -0.0019 0.0648 1.0000
15.000 1.0663 0.07392 0.06707 -0.0026 0.0600 1.0000
15.250 1.0653 0.07737 0.07060 -0.0034 0.0558 1.0000
15.500 1.0633 0.08104 0.07431 -0.0044 0.0522 1.0000
15.750 1.0630 0.08456 0.07795 -0.0054 0.0485 1.0000
16.000 1.0606 0.08841 0.08181 -0.0067 0.0458 1.0000
16.250 1.0602 0.09210 0.08565 -0.0079 0.0425 1.0000
16.500 1.0583 0.09602 0.08962 -0.0093 0.0403 1.0000
16.750 1.0575 0.09989 0.09365 -0.0107 0.0377 1.0000
17.000 1.0551 0.10400 0.09777 -0.0125 0.0357 1.0000
17.250 1.0537 0.10811 0.10208 -0.0142 0.0334 1.0000
17.500 1.0521 0.11222 0.10623 -0.0161 0.0319 1.0000
17.750 1.0504 0.11645 0.11062 -0.0180 0.0301 1.0000
18.000 1.0476 0.12095 0.11523 -0.0202 0.0285 1.0000
18.250 1.0458 0.12524 0.11957 -0.0225 0.0272 1.0000
18.500 1.0419 0.13013 0.12466 -0.0250 0.0259 1.0000
18.750 1.0392 0.13475 0.12939 -0.0276 0.0248 1.0000
19.000 1.0369 0.13928 0.13397 -0.0303 0.0238 1.0000
19.250 1.0286 0.14538 0.14030 -0.0338 0.0229 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 545 AIRFOIL (e545-il)