EPPLER 520 AIRFOIL (e520-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 520 AIRFOIL (e520-il) Reynolds number: 100,000 Max Cl/Cd: 41.21 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e520-il-100000.txt Download as CSV file: xf-e520-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 520 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
------ -------- --------- --------- -------- -------- --------
-15.000 -0.6712 0.14996 0.14512 0.0059 1.0000 0.0768
-14.750 -0.6569 0.15030 0.14544 0.0066 1.0000 0.0894
-14.500 -0.8225 0.09832 0.09337 -0.0245 1.0000 0.0528
-14.250 -0.8489 0.09032 0.08526 -0.0290 1.0000 0.0524
-14.000 -0.8737 0.08318 0.07797 -0.0326 1.0000 0.0518
-13.750 -0.8987 0.07684 0.07145 -0.0353 1.0000 0.0514
-13.500 -0.9212 0.07128 0.06569 -0.0371 1.0000 0.0509
-13.250 -0.9452 0.06647 0.06062 -0.0381 1.0000 0.0512
-13.000 -0.9640 0.06219 0.05609 -0.0382 1.0000 0.0513
-12.750 -0.9827 0.05833 0.05194 -0.0376 1.0000 0.0516
-12.500 -0.9987 0.05509 0.04839 -0.0364 1.0000 0.0522
-12.250 -1.0137 0.05240 0.04536 -0.0344 1.0000 0.0528
-12.000 -1.0057 0.04798 0.04085 -0.0340 1.0000 0.0553
-11.750 -1.0005 0.04607 0.03889 -0.0328 1.0000 0.0583
-11.500 -0.9985 0.04392 0.03653 -0.0311 1.0000 0.0613
-11.250 -0.9994 0.04190 0.03413 -0.0285 1.0000 0.0640
-11.000 -0.9775 0.03898 0.03124 -0.0288 1.0000 0.0692
-10.750 -0.9686 0.03738 0.02944 -0.0271 1.0000 0.0744
-10.500 -0.9467 0.03498 0.02697 -0.0270 1.0000 0.0811
-10.250 -0.9347 0.03358 0.02544 -0.0256 1.0000 0.0884
-10.000 -0.9131 0.03171 0.02367 -0.0254 1.0000 0.0980
-9.750 -0.8963 0.03012 0.02209 -0.0246 1.0000 0.1090
-9.500 -0.8813 0.02870 0.02071 -0.0235 1.0000 0.1218
-9.250 -0.8677 0.02739 0.01948 -0.0223 1.0000 0.1367
-9.000 -0.8568 0.02624 0.01843 -0.0208 1.0000 0.1533
-8.750 -0.8524 0.02535 0.01760 -0.0181 1.0000 0.1697
-8.500 -0.8613 0.02486 0.01720 -0.0130 1.0000 0.1812
-8.250 -0.8701 0.02435 0.01674 -0.0078 1.0000 0.1939
-8.000 -0.8744 0.02372 0.01619 -0.0034 1.0000 0.2101
-7.750 -0.8755 0.02305 0.01556 0.0006 1.0000 0.2309
-7.500 -0.8738 0.02229 0.01497 0.0041 1.0000 0.2538
-7.250 -0.8693 0.02154 0.01437 0.0071 0.9998 0.2815
-7.000 -0.8440 0.02065 0.01370 0.0062 0.9942 0.3265
-6.750 -0.8202 0.01991 0.01319 0.0058 0.9885 0.3732
-6.500 -0.7939 0.01938 0.01282 0.0053 0.9834 0.4234
-6.250 -0.7686 0.01909 0.01276 0.0053 0.9785 0.4702
-6.000 -0.7420 0.01909 0.01290 0.0054 0.9732 0.5134
-5.750 -0.7084 0.01951 0.01341 0.0047 0.9690 0.5520
-5.500 -0.6814 0.02008 0.01395 0.0053 0.9642 0.5810
-5.250 -0.6515 0.02083 0.01466 0.0055 0.9592 0.6051
-5.000 -0.6159 0.02166 0.01538 0.0047 0.9550 0.6275
-4.750 -0.5884 0.02245 0.01609 0.0054 0.9504 0.6450
-4.500 -0.5604 0.02327 0.01682 0.0061 0.9453 0.6601
-4.250 -0.5240 0.02410 0.01756 0.0053 0.9412 0.6747
-4.000 -0.4939 0.02477 0.01814 0.0055 0.9368 0.6888
-3.750 -0.4686 0.02545 0.01875 0.0067 0.9313 0.7001
-3.500 -0.4299 0.02621 0.01944 0.0058 0.9273 0.7107
-3.250 -0.3849 0.02678 0.01991 0.0033 0.9241 0.7225
-3.000 -0.3792 0.02691 0.01999 0.0072 0.9166 0.7350
-2.750 -0.3444 0.02719 0.02018 0.0061 0.9120 0.7478
-2.500 -0.2855 0.02790 0.02082 0.0018 0.9097 0.7548
-2.250 -0.2851 0.02795 0.02084 0.0064 0.9013 0.7663
-2.000 -0.2473 0.02806 0.02088 0.0045 0.8968 0.7780
-1.750 -0.2052 0.02847 0.02125 0.0027 0.8922 0.7853
-1.500 -0.1878 0.02851 0.02125 0.0043 0.8849 0.7956
-1.250 -0.1495 0.02846 0.02115 0.0021 0.8805 0.8066
-1.000 -0.1247 0.02877 0.02146 0.0030 0.8728 0.8132
-0.750 -0.0912 0.02867 0.02132 0.0017 0.8672 0.8228
-0.500 -0.0614 0.02887 0.02152 0.0014 0.8598 0.8300
-0.250 -0.0306 0.02877 0.02140 0.0006 0.8533 0.8385
0.000 -0.0001 0.02888 0.02151 0.0000 0.8456 0.8456
0.250 0.0309 0.02877 0.02139 -0.0007 0.8384 0.8532
0.500 0.0615 0.02886 0.02151 -0.0014 0.8299 0.8598
0.750 0.0913 0.02867 0.02132 -0.0017 0.8229 0.8672
1.000 0.1244 0.02877 0.02146 -0.0029 0.8133 0.8728
1.250 0.1491 0.02848 0.02117 -0.0021 0.8067 0.8806
1.500 0.1869 0.02853 0.02128 -0.0042 0.7958 0.8850
1.750 0.2045 0.02848 0.02126 -0.0026 0.7854 0.8923
2.000 0.2469 0.02806 0.02088 -0.0045 0.7780 0.8969
2.250 0.2848 0.02795 0.02084 -0.0064 0.7663 0.9013
2.500 0.2858 0.02787 0.02078 -0.0018 0.7546 0.9097
2.750 0.3448 0.02716 0.02015 -0.0061 0.7478 0.9120
3.000 0.3789 0.02691 0.01999 -0.0071 0.7351 0.9167
3.250 0.3846 0.02679 0.01992 -0.0033 0.7226 0.9242
3.500 0.4298 0.02621 0.01944 -0.0058 0.7108 0.9273
3.750 0.4687 0.02543 0.01872 -0.0067 0.7005 0.9313
4.000 0.4930 0.02478 0.01815 -0.0054 0.6888 0.9369
4.250 0.5239 0.02410 0.01756 -0.0053 0.6748 0.9412
4.500 0.5601 0.02326 0.01681 -0.0061 0.6601 0.9453
4.750 0.5878 0.02245 0.01609 -0.0053 0.6451 0.9505
5.000 0.6159 0.02164 0.01536 -0.0047 0.6279 0.9551
5.250 0.6512 0.02083 0.01466 -0.0055 0.6052 0.9593
5.500 0.6810 0.02007 0.01394 -0.0052 0.5812 0.9643
5.750 0.7083 0.01950 0.01339 -0.0047 0.5520 0.9691
6.000 0.7419 0.01907 0.01288 -0.0054 0.5134 0.9733
6.250 0.7683 0.01908 0.01274 -0.0052 0.4701 0.9787
6.500 0.7938 0.01937 0.01281 -0.0052 0.4230 0.9835
6.750 0.8201 0.01990 0.01318 -0.0058 0.3731 0.9887
7.000 0.8440 0.02063 0.01369 -0.0062 0.3266 0.9943
7.250 0.8693 0.02153 0.01435 -0.0071 0.2816 0.9999
7.500 0.8735 0.02228 0.01496 -0.0040 0.2541 1.0000
7.750 0.8752 0.02304 0.01553 -0.0005 0.2313 1.0000
8.000 0.8742 0.02371 0.01618 0.0034 0.2104 1.0000
8.250 0.8699 0.02433 0.01672 0.0078 0.1942 1.0000
8.500 0.8609 0.02483 0.01717 0.0130 0.1815 1.0000
8.750 0.8525 0.02534 0.01758 0.0181 0.1699 1.0000
9.000 0.8574 0.02625 0.01844 0.0206 0.1533 1.0000
9.250 0.8683 0.02740 0.01949 0.0222 0.1365 1.0000
9.500 0.8821 0.02870 0.02071 0.0234 0.1217 1.0000
9.750 0.8965 0.03010 0.02206 0.0245 0.1088 1.0000
10.000 0.9141 0.03173 0.02368 0.0252 0.0979 1.0000
10.250 0.9359 0.03361 0.02548 0.0254 0.0884 1.0000
10.500 0.9474 0.03500 0.02697 0.0268 0.0808 1.0000
10.750 0.9696 0.03741 0.02948 0.0269 0.0743 1.0000
11.000 0.9789 0.03903 0.03130 0.0285 0.0691 1.0000
11.250 0.9994 0.04188 0.03413 0.0284 0.0639 1.0000
11.500 0.9994 0.04399 0.03660 0.0309 0.0612 1.0000
11.750 1.0013 0.04594 0.03875 0.0327 0.0578 1.0000
12.000 1.0091 0.04795 0.04075 0.0336 0.0547 1.0000
12.250 1.0132 0.05238 0.04536 0.0343 0.0527 1.0000
12.500 0.9995 0.05521 0.04852 0.0362 0.0522 1.0000
12.750 0.9830 0.05846 0.05209 0.0375 0.0516 1.0000
13.000 0.9648 0.06228 0.05619 0.0380 0.0512 1.0000
13.250 0.9450 0.06654 0.06070 0.0378 0.0511 1.0000
13.500 0.9239 0.07145 0.06585 0.0369 0.0513 1.0000
13.750 0.8996 0.07703 0.07164 0.0350 0.0514 1.0000
14.000 0.8738 0.08331 0.07811 0.0322 0.0517 1.0000
14.250 0.8468 0.09059 0.08554 0.0283 0.0521 1.0000
14.500 0.8232 0.09854 0.09359 0.0242 0.0528 1.0000
14.750 0.7005 0.13368 0.12894 0.0013 0.0657 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 520 AIRFOIL (e520-il)