EPPLER 560 AIRFOIL (e560-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 560 AIRFOIL (e560-il) Reynolds number: 50,000 Max Cl/Cd: 7.82 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e560-il-50000.txt Download as CSV file: xf-e560-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 560 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.2734 0.13093 0.12394 -0.0338 1.0000 0.2505
-10.000 -0.2528 0.12558 0.11859 -0.0325 1.0000 0.2586
-9.750 -0.2687 0.12516 0.11829 -0.0318 1.0000 0.2664
-9.500 -0.2543 0.12071 0.11385 -0.0304 1.0000 0.2752
-9.250 -0.2729 0.12052 0.11379 -0.0292 1.0000 0.2829
-9.000 -0.2589 0.11638 0.10966 -0.0275 1.0000 0.2944
-8.750 -0.2851 0.11683 0.11027 -0.0257 1.0000 0.3000
-8.500 -0.2694 0.11270 0.10614 -0.0238 1.0000 0.3128
-8.250 -0.3022 0.11368 0.10731 -0.0211 1.0000 0.3171
-8.000 -0.2849 0.10959 0.10321 -0.0191 1.0000 0.3305
-7.750 -0.3196 0.11066 0.10447 -0.0159 1.0000 0.3340
-7.500 -0.3022 0.10676 0.10054 -0.0139 1.0000 0.3470
-7.250 -0.3374 0.10775 0.10172 -0.0104 1.0000 0.3509
-7.000 -0.3219 0.10403 0.09797 -0.0086 1.0000 0.3608
-6.750 -0.3511 0.10450 0.09858 -0.0052 1.0000 0.3673
-6.500 -0.3493 0.10194 0.09606 -0.0032 1.0000 0.3750
-6.250 -0.3694 0.10171 0.09592 0.0000 1.0000 0.3837
-6.000 -0.4971 0.09111 0.08564 -0.0115 1.0000 0.2674
-5.750 -0.4907 0.09112 0.08566 -0.0069 1.0000 0.2828
-5.500 -0.4968 0.08757 0.08210 -0.0073 1.0000 0.2810
-5.250 -0.5031 0.08316 0.07769 -0.0098 0.9997 0.2789
-5.000 -0.4432 0.08788 0.08230 -0.0043 0.9869 0.3426
-4.750 -0.4412 0.07090 0.06497 -0.0329 0.9775 0.2972
-4.500 -0.4080 0.06852 0.06243 -0.0375 0.9661 0.3113
-4.250 -0.3737 0.06528 0.05898 -0.0441 0.9547 0.3251
-4.000 -0.3346 0.06161 0.05504 -0.0523 0.9438 0.3394
-3.750 -0.2915 0.05909 0.05226 -0.0592 0.9330 0.3557
-3.500 -0.2527 0.05644 0.04932 -0.0659 0.9212 0.3710
-3.250 -0.2129 0.05431 0.04689 -0.0723 0.9099 0.3867
-3.000 -0.1883 0.05450 0.04710 -0.0707 0.8988 0.3974
-2.750 -0.1448 0.05317 0.04558 -0.0756 0.8883 0.4126
-2.500 -0.1137 0.05195 0.04415 -0.0791 0.8767 0.4257
-2.250 -0.0733 0.05088 0.04284 -0.0839 0.8660 0.4407
-2.000 -0.0278 0.05001 0.04175 -0.0886 0.8559 0.4568
-1.750 -0.0121 0.05006 0.04182 -0.0871 0.8445 0.4657
-1.500 0.0272 0.04948 0.04107 -0.0906 0.8346 0.4795
-1.250 0.0642 0.04902 0.04044 -0.0937 0.8242 0.4939
-1.000 0.0903 0.04899 0.04029 -0.0953 0.8135 0.5061
-0.750 0.1359 0.04856 0.03976 -0.0984 0.8051 0.5216
-0.500 0.1492 0.04896 0.04011 -0.0978 0.7937 0.5310
-0.250 0.1858 0.04891 0.03994 -0.1003 0.7847 0.5457
0.000 0.2143 0.04915 0.04007 -0.1020 0.7748 0.5597
0.250 0.2325 0.04965 0.04059 -0.1013 0.7654 0.5704
0.500 0.2650 0.04981 0.04070 -0.1028 0.7568 0.5852
0.750 0.2816 0.05067 0.04150 -0.1029 0.7473 0.5975
1.000 0.3182 0.05078 0.04157 -0.1047 0.7393 0.6142
1.250 0.3277 0.05206 0.04283 -0.1041 0.7300 0.6266
1.500 0.3638 0.05231 0.04303 -0.1057 0.7224 0.6449
1.750 0.3678 0.05381 0.04457 -0.1041 0.7138 0.6563
2.000 0.3975 0.05437 0.04511 -0.1050 0.7062 0.6755
2.250 0.4074 0.05585 0.04661 -0.1043 0.6987 0.6907
2.500 0.4236 0.05704 0.04782 -0.1042 0.6912 0.7092
2.750 0.4500 0.05782 0.04862 -0.1045 0.6844 0.7329
3.000 0.4460 0.06000 0.05087 -0.1030 0.6779 0.7495
3.250 0.4737 0.06055 0.05150 -0.1030 0.6716 0.7816
3.500 0.4729 0.06236 0.05343 -0.1011 0.6659 0.8084
3.750 0.4689 0.06419 0.05544 -0.0990 0.6615 0.8440
4.000 0.4759 0.06524 0.05673 -0.0982 0.6572 1.0000
4.250 0.5212 0.06768 0.05891 -0.1054 0.6495 1.0000
4.500 0.5301 0.07086 0.06195 -0.1078 0.6466 1.0000
4.750 0.5393 0.07393 0.06489 -0.1094 0.6455 1.0000
5.000 0.5463 0.07703 0.06790 -0.1105 0.6458 1.0000
5.250 0.5569 0.08032 0.07111 -0.1119 0.6496 1.0000
5.500 0.4827 0.08970 0.08074 -0.1136 0.7539 1.0000
5.750 0.4829 0.09069 0.08168 -0.1124 0.7419 1.0000
6.000 0.5109 0.09407 0.08495 -0.1146 0.7362 1.0000
6.250 0.5209 0.09566 0.08650 -0.1143 0.7237 1.0000
6.500 0.5312 0.09793 0.08874 -0.1144 0.7144 1.0000
6.750 0.5619 0.10106 0.09180 -0.1164 0.7049 1.0000
7.000 0.5624 0.10274 0.09348 -0.1155 0.6944 1.0000
7.250 0.6023 0.10689 0.09759 -0.1182 0.6867 1.0000
7.500 0.5973 0.10789 0.09860 -0.1168 0.6741 1.0000
7.750 0.6172 0.11122 0.10191 -0.1179 0.6677 1.0000
8.000 0.6314 0.11337 0.10407 -0.1181 0.6554 1.0000
8.250 0.6356 0.11558 0.10630 -0.1178 0.6450 1.0000
8.500 0.6690 0.11943 0.11016 -0.1196 0.6368 1.0000
8.750 0.6639 0.12091 0.11167 -0.1187 0.6253 1.0000
9.000 0.6956 0.12520 0.11597 -0.1205 0.6184 1.0000
9.250 0.6961 0.12661 0.11741 -0.1199 0.6055 1.0000
9.500 0.7025 0.12936 0.12021 -0.1201 0.5973 1.0000
9.750 0.7287 0.13277 0.12366 -0.1212 0.5869 1.0000
10.000 0.7257 0.13465 0.12558 -0.1209 0.5760 1.0000
10.250 0.7636 0.13963 0.13060 -0.1227 0.5688 1.0000
10.500 0.7518 0.14043 0.13145 -0.1220 0.5568 1.0000
10.750 0.7666 0.14394 0.13501 -0.1227 0.5491 1.0000
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Polar data table (+)
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