EPPLER 552 AIRFOIL (e552-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: EPPLER 552 AIRFOIL (e552-il) Reynolds number: 50,000 Max Cl/Cd: 10.95 at α=-3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e552-il-50000.txt Download as CSV file: xf-e552-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 552 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.4416 0.09972 0.09310 -0.0574 1.0000 0.1155
-11.500 -0.4547 0.09290 0.08631 -0.0602 1.0000 0.1129
-11.250 -0.4731 0.08602 0.07949 -0.0630 1.0000 0.1101
-11.000 -0.6013 0.07458 0.06784 -0.0719 1.0000 0.1003
-10.750 -0.6143 0.07078 0.06405 -0.0713 1.0000 0.0996
-10.500 -0.6365 0.06773 0.06100 -0.0695 1.0000 0.0990
-10.250 -0.6558 0.06484 0.05810 -0.0673 1.0000 0.0983
-10.000 -0.6765 0.06220 0.05540 -0.0645 1.0000 0.0979
-9.750 -0.6961 0.06015 0.05332 -0.0610 1.0000 0.0975
-9.500 -0.7184 0.05873 0.05186 -0.0566 1.0000 0.0971
-9.250 -0.7406 0.05752 0.05058 -0.0522 1.0000 0.0968
-9.000 -0.7576 0.05613 0.04907 -0.0486 1.0000 0.0965
-8.750 -0.7700 0.05435 0.04709 -0.0456 1.0000 0.0966
-8.500 -0.7761 0.05243 0.04493 -0.0431 1.0000 0.0968
-8.250 -0.7789 0.05045 0.04261 -0.0410 1.0000 0.0982
-8.000 -0.7639 0.04799 0.03964 -0.0417 0.9957 0.1009
-7.750 -0.7207 0.04502 0.03674 -0.0452 0.9860 0.1095
-7.500 -0.6788 0.04251 0.03407 -0.0479 0.9760 0.1200
-7.250 -0.6374 0.04056 0.03219 -0.0498 0.9661 0.1380
-7.000 -0.5993 0.03863 0.03053 -0.0514 0.9569 0.1672
-6.750 -0.5790 0.03627 0.02868 -0.0518 0.9460 0.2104
-6.500 -0.5749 0.03386 0.02802 -0.0501 0.9349 0.3178
-6.250 -0.5193 0.04673 0.04114 -0.0379 0.9216 0.4942
-6.000 -0.3590 0.06086 0.05449 -0.0309 0.9171 0.5635
-5.750 -0.1776 0.06318 0.05596 -0.0409 0.9167 0.6546
-5.000 0.2795 0.04543 0.03668 -0.0989 0.9137 1.0000
-4.750 0.3101 0.04413 0.03522 -0.1030 0.9019 1.0000
-4.500 0.3409 0.04294 0.03388 -0.1068 0.8916 1.0000
-4.250 0.3709 0.04187 0.03265 -0.1102 0.8818 1.0000
-4.000 0.3887 0.04152 0.03222 -0.1112 0.8706 1.0000
-3.750 0.4127 0.04094 0.03153 -0.1131 0.8621 1.0000
-3.500 0.4280 0.04091 0.03143 -0.1135 0.8526 1.0000
-3.250 0.4465 0.04077 0.03121 -0.1143 0.8446 1.0000
-3.000 0.1970 0.05163 0.04246 -0.0697 0.8324 0.8499
-2.750 0.0924 0.05457 0.04559 -0.0526 0.8220 0.8132
-2.500 0.0362 0.05550 0.04658 -0.0425 0.8165 0.8003
-2.250 -0.1202 0.05721 0.04861 -0.0174 0.8137 0.7851
-2.000 -0.1725 0.05710 0.04859 -0.0080 0.8117 0.7836
-1.750 -0.2096 0.05660 0.04811 -0.0013 0.8108 0.7840
-1.500 -0.2354 0.05596 0.04745 0.0034 0.8110 0.7857
-1.250 -0.2400 0.05560 0.04704 0.0056 0.8114 0.7883
-1.000 -0.2394 0.05534 0.04670 0.0070 0.8119 0.7913
-0.750 -0.2369 0.05511 0.04641 0.0080 0.8128 0.7946
-0.500 -0.2282 0.05503 0.04624 0.0077 0.8147 0.7981
-0.250 -0.2475 0.05474 0.04594 0.0104 0.8282 0.8007
0.000 -0.4450 0.04839 0.04016 0.0409 1.0000 0.7996
0.250 -0.4237 0.04838 0.04002 0.0387 1.0000 0.8032
0.500 -0.4055 0.04841 0.03995 0.0379 1.0000 0.8071
0.750 -0.3850 0.04860 0.04002 0.0364 0.9998 0.8115
1.000 -0.3482 0.04996 0.04122 0.0312 0.9938 0.8159
1.250 -0.3181 0.05083 0.04199 0.0278 0.9866 0.8202
1.500 -0.2886 0.05190 0.04295 0.0247 0.9787 0.8251
1.750 -0.2530 0.05349 0.04442 0.0202 0.9703 0.8305
2.000 -0.2190 0.05478 0.04562 0.0160 0.9592 0.8352
2.250 -0.1940 0.05551 0.04629 0.0140 0.9479 0.8404
2.500 -0.1660 0.05664 0.04735 0.0111 0.9374 0.8467
2.750 -0.1301 0.05876 0.04940 0.0071 0.9279 0.8532
3.000 -0.1052 0.05947 0.05007 0.0051 0.9147 0.8597
3.250 -0.0818 0.06033 0.05090 0.0031 0.9026 0.8663
3.500 -0.0568 0.06165 0.05220 0.0013 0.8920 0.8741
3.750 -0.0222 0.06375 0.05428 -0.0022 0.8811 0.8837
4.000 -0.0041 0.06424 0.05478 -0.0030 0.8678 0.8930
4.250 0.0172 0.06530 0.05587 -0.0043 0.8558 0.9029
4.500 0.0509 0.06760 0.05820 -0.0078 0.8467 0.9164
4.750 0.0803 0.06905 0.05971 -0.0107 0.8330 0.9315
5.000 0.1065 0.07022 0.06096 -0.0136 0.8194 0.9492
5.250 0.1380 0.07187 0.06269 -0.0179 0.8057 0.9763
5.500 0.1656 0.07374 0.06455 -0.0217 0.7940 1.0000
5.750 0.2116 0.07729 0.06808 -0.0283 0.7833 1.0000
6.000 0.2301 0.07841 0.06918 -0.0306 0.7694 1.0000
6.250 0.2524 0.08026 0.07102 -0.0335 0.7565 1.0000
6.500 0.2806 0.08283 0.07357 -0.0374 0.7457 1.0000
6.750 0.3249 0.08668 0.07741 -0.0431 0.7346 1.0000
7.000 0.3394 0.08804 0.07876 -0.0447 0.7205 1.0000
7.250 0.3556 0.09000 0.08071 -0.0465 0.7080 1.0000
7.500 0.3789 0.09276 0.08346 -0.0490 0.6974 1.0000
7.750 0.4171 0.09657 0.08725 -0.0529 0.6860 1.0000
8.000 0.4239 0.09786 0.08855 -0.0531 0.6724 1.0000
8.250 0.4352 0.09997 0.09066 -0.0539 0.6601 1.0000
8.500 0.4579 0.10314 0.09385 -0.0557 0.6505 1.0000
8.750 0.4855 0.10632 0.09705 -0.0577 0.6380 1.0000
9.000 0.4893 0.10789 0.09864 -0.0577 0.6247 1.0000
9.250 0.4985 0.11030 0.10108 -0.0582 0.6137 1.0000
9.500 0.5299 0.11442 0.10525 -0.0604 0.6042 1.0000
9.750 0.5397 0.11637 0.10726 -0.0608 0.5907 1.0000
10.000 0.5416 0.11834 0.10926 -0.0609 0.5787 1.0000
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Polar data table (+)
Polar graphs
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