EPPLER E852 AIRFOIL (e852-il) Xfoil prediction polar at RE=500,000 Ncrit=5
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Airfoil: EPPLER E852 AIRFOIL (e852-il) Reynolds number: 500,000 Max Cl/Cd: 98.95 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e852-il-500000-n5.txt Download as CSV file: xf-e852-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER E852 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4462 0.09248 0.09027 -0.0405 1.0000 0.0026
-9.500 -0.4527 0.08793 0.08576 -0.0416 1.0000 0.0025
-9.250 -0.4595 0.08345 0.08133 -0.0429 0.9999 0.0026
-9.000 -0.4597 0.07500 0.07291 -0.0507 0.9975 0.0025
-8.750 -0.4611 0.06403 0.06196 -0.0628 0.9931 0.0024
-8.500 -0.4692 0.04974 0.04746 -0.0829 0.9819 0.0024
-8.250 -0.4701 0.04205 0.03946 -0.0887 0.9702 0.0025
-7.750 -0.4543 0.02882 0.02532 -0.0933 0.9492 0.0027
-7.500 -0.4299 0.02488 0.02091 -0.0955 0.9443 0.0029
-7.250 -0.4036 0.02207 0.01771 -0.0969 0.9385 0.0031
-7.000 -0.3730 0.01956 0.01482 -0.0988 0.9338 0.0033
-6.750 -0.3386 0.01848 0.01353 -0.1009 0.9302 0.0039
-6.500 -0.3074 0.01733 0.01217 -0.1022 0.9234 0.0041
-6.250 -0.2746 0.01581 0.01035 -0.1038 0.9177 0.0044
-6.000 -0.2461 0.01412 0.00840 -0.1045 0.9097 0.0045
-5.750 -0.2167 0.01285 0.00692 -0.1053 0.9021 0.0046
-5.500 -0.1892 0.01178 0.00567 -0.1057 0.8934 0.0048
-5.250 -0.1618 0.01099 0.00471 -0.1059 0.8850 0.0051
-4.750 -0.1063 0.01003 0.00348 -0.1062 0.8698 0.0078
-4.500 -0.0785 0.00965 0.00304 -0.1063 0.8630 0.0183
-4.250 -0.0517 0.00930 0.00275 -0.1065 0.8559 0.0405
-4.000 -0.0243 0.00894 0.00245 -0.1067 0.8499 0.0748
-3.750 0.0024 0.00856 0.00220 -0.1068 0.8436 0.1241
-3.500 0.0294 0.00814 0.00196 -0.1071 0.8379 0.1932
-3.250 0.0554 0.00751 0.00173 -0.1074 0.8321 0.3168
-3.000 0.0826 0.00723 0.00163 -0.1075 0.8268 0.3927
-2.750 0.1102 0.00710 0.00157 -0.1076 0.8219 0.4388
-2.500 0.1379 0.00703 0.00151 -0.1077 0.8167 0.4652
-2.250 0.1659 0.00701 0.00145 -0.1078 0.8121 0.4835
-2.000 0.1937 0.00697 0.00143 -0.1079 0.8077 0.5050
-1.750 0.2214 0.00694 0.00143 -0.1079 0.8029 0.5265
-1.500 0.2494 0.00694 0.00140 -0.1080 0.7987 0.5401
-1.250 0.2773 0.00693 0.00137 -0.1081 0.7946 0.5498
-1.000 0.3052 0.00692 0.00137 -0.1082 0.7901 0.5573
-0.500 0.3612 0.00694 0.00137 -0.1084 0.7824 0.5723
-0.250 0.3891 0.00694 0.00139 -0.1085 0.7780 0.5799
0.000 0.4170 0.00695 0.00142 -0.1085 0.7739 0.5878
0.250 0.4449 0.00697 0.00144 -0.1086 0.7702 0.5956
0.500 0.4726 0.00698 0.00149 -0.1086 0.7652 0.6043
0.750 0.5001 0.00699 0.00153 -0.1086 0.7600 0.6129
1.000 0.5275 0.00701 0.00160 -0.1086 0.7540 0.6220
1.250 0.5547 0.00703 0.00164 -0.1085 0.7464 0.6319
1.500 0.5815 0.00704 0.00170 -0.1083 0.7372 0.6420
1.750 0.6080 0.00707 0.00176 -0.1080 0.7269 0.6527
2.000 0.6342 0.00710 0.00185 -0.1077 0.7143 0.6640
2.250 0.6597 0.00714 0.00191 -0.1071 0.6967 0.6763
2.500 0.6845 0.00720 0.00198 -0.1065 0.6744 0.6895
2.750 0.7092 0.00729 0.00207 -0.1058 0.6510 0.7035
3.000 0.7332 0.00741 0.00219 -0.1050 0.6208 0.7190
3.250 0.7523 0.00777 0.00235 -0.1032 0.5544 0.7359
3.500 0.7662 0.00854 0.00273 -0.1006 0.4504 0.7540
3.750 0.7805 0.00940 0.00319 -0.0982 0.3461 0.7747
4.000 0.7958 0.01021 0.00365 -0.0961 0.2504 0.7988
4.250 0.8102 0.01106 0.00414 -0.0938 0.1565 0.8278
4.500 0.8231 0.01186 0.00463 -0.0911 0.0767 0.8676
4.750 0.8408 0.01222 0.00496 -0.0891 0.0433 0.9645
5.000 0.8625 0.01279 0.00538 -0.0882 0.0191 1.0000
5.250 0.8838 0.01336 0.00586 -0.0872 0.0050 1.0000
5.500 0.9058 0.01386 0.00638 -0.0862 0.0025 1.0000
5.750 0.9276 0.01438 0.00698 -0.0851 0.0020 1.0000
6.000 0.9489 0.01495 0.00771 -0.0840 0.0018 1.0000
6.250 0.9681 0.01574 0.00863 -0.0824 0.0017 1.0000
6.500 0.9873 0.01650 0.00949 -0.0809 0.0016 1.0000
6.750 1.0058 0.01733 0.01043 -0.0794 0.0016 1.0000
7.000 1.0235 0.01828 0.01148 -0.0776 0.0016 1.0000
7.250 1.0403 0.01938 0.01272 -0.0758 0.0016 1.0000
7.500 1.0573 0.02059 0.01406 -0.0740 0.0015 1.0000
7.750 1.0746 0.02197 0.01559 -0.0724 0.0015 1.0000
8.000 1.0924 0.02360 0.01741 -0.0708 0.0015 1.0000
8.250 1.1105 0.02541 0.01944 -0.0694 0.0015 1.0000
8.500 1.1274 0.02753 0.02182 -0.0678 0.0015 1.0000
8.750 1.1420 0.02986 0.02444 -0.0659 0.0016 1.0000
9.000 1.1525 0.03255 0.02745 -0.0635 0.0016 1.0000
9.250 1.1569 0.03549 0.03074 -0.0603 0.0016 1.0000
9.500 1.1556 0.03843 0.03406 -0.0564 0.0016 1.0000
9.750 1.1483 0.04162 0.03757 -0.0520 0.0016 1.0000
10.000 1.1372 0.04493 0.04116 -0.0477 0.0016 1.0000
10.250 1.1211 0.04864 0.04515 -0.0437 0.0016 1.0000
10.750 1.0866 0.05645 0.05342 -0.0383 0.0017 1.0000
11.000 1.0724 0.06039 0.05754 -0.0371 0.0017 1.0000
11.250 1.0571 0.06477 0.06211 -0.0367 0.0017 1.0000
11.500 1.0415 0.06964 0.06713 -0.0373 0.0017 1.0000
11.750 1.0234 0.07546 0.07312 -0.0390 0.0017 1.0000
12.000 1.0052 0.08211 0.07992 -0.0421 0.0017 1.0000
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Polar data table (+)
Polar graphs
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