E221 (9.39%) (e221-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: E221 (9.39%) (e221-il) Reynolds number: 500,000 Max Cl/Cd: 86.66 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e221-il-500000.txt Download as CSV file: xf-e221-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: E221 (9.39%)
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5336 0.08144 0.07935 -0.0104 1.0000 0.0188
-8.500 -0.5394 0.07516 0.07311 -0.0160 1.0000 0.0189
-8.250 -0.5505 0.06838 0.06631 -0.0217 1.0000 0.0189
-8.000 -0.5585 0.06239 0.06019 -0.0243 1.0000 0.0189
-7.750 -0.5654 0.05694 0.05454 -0.0256 1.0000 0.0189
-5.500 -0.4676 0.02073 0.01577 -0.0193 0.9793 0.0179
-5.250 -0.4362 0.01626 0.01055 -0.0192 0.9595 0.0144
-5.000 -0.4057 0.01508 0.00918 -0.0195 0.9335 0.0142
-4.750 -0.3823 0.01389 0.00779 -0.0182 0.9055 0.0142
-4.500 -0.3611 0.01282 0.00655 -0.0165 0.8810 0.0146
-4.250 -0.3387 0.01206 0.00567 -0.0151 0.8597 0.0152
-4.000 -0.3150 0.01156 0.00505 -0.0141 0.8416 0.0164
-3.750 -0.2905 0.01110 0.00447 -0.0132 0.8249 0.0175
-3.500 -0.2651 0.01075 0.00400 -0.0124 0.8094 0.0188
-3.250 -0.2404 0.01025 0.00342 -0.0117 0.7952 0.0229
-3.000 -0.2152 0.00974 0.00287 -0.0108 0.7818 0.0384
-2.750 -0.1911 0.00915 0.00253 -0.0101 0.7691 0.1070
-2.500 -0.1659 0.00877 0.00233 -0.0097 0.7564 0.1691
-2.250 -0.1414 0.00830 0.00217 -0.0091 0.7444 0.2589
-2.000 -0.1182 0.00772 0.00202 -0.0084 0.7328 0.3881
-1.750 -0.0962 0.00714 0.00190 -0.0073 0.7218 0.5269
-1.500 -0.0752 0.00663 0.00180 -0.0058 0.7109 0.6564
-1.250 -0.0551 0.00609 0.00176 -0.0038 0.7000 0.7978
-1.000 -0.0014 0.00601 0.00194 -0.0085 0.6885 0.9267
-0.750 0.0372 0.00620 0.00202 -0.0104 0.6774 0.9558
-0.500 0.0754 0.00639 0.00208 -0.0123 0.6662 0.9708
-0.250 0.1238 0.00651 0.00209 -0.0165 0.6536 0.9796
0.000 0.1684 0.00660 0.00208 -0.0200 0.6413 0.9871
0.250 0.2129 0.00666 0.00204 -0.0235 0.6285 0.9937
0.500 0.2588 0.00664 0.00193 -0.0274 0.6157 0.9991
0.750 0.2875 0.00665 0.00187 -0.0277 0.6038 1.0000
1.000 0.3124 0.00669 0.00183 -0.0271 0.5921 1.0000
1.250 0.3374 0.00672 0.00180 -0.0265 0.5797 1.0000
1.500 0.3626 0.00676 0.00179 -0.0260 0.5672 1.0000
1.750 0.3879 0.00681 0.00178 -0.0255 0.5546 1.0000
2.000 0.4133 0.00686 0.00180 -0.0250 0.5419 1.0000
2.250 0.4387 0.00693 0.00181 -0.0245 0.5289 1.0000
2.500 0.4642 0.00700 0.00184 -0.0240 0.5158 1.0000
2.750 0.4896 0.00709 0.00188 -0.0235 0.5027 1.0000
3.000 0.5151 0.00718 0.00193 -0.0230 0.4894 1.0000
3.250 0.5405 0.00729 0.00201 -0.0225 0.4755 1.0000
3.500 0.5659 0.00740 0.00209 -0.0220 0.4614 1.0000
3.750 0.5912 0.00752 0.00218 -0.0215 0.4469 1.0000
4.000 0.6165 0.00765 0.00229 -0.0210 0.4316 1.0000
4.250 0.6417 0.00780 0.00241 -0.0205 0.4157 1.0000
4.500 0.6668 0.00795 0.00254 -0.0200 0.3995 1.0000
4.750 0.6918 0.00813 0.00268 -0.0195 0.3816 1.0000
5.000 0.7164 0.00834 0.00285 -0.0189 0.3617 1.0000
5.250 0.7411 0.00856 0.00303 -0.0184 0.3399 1.0000
5.500 0.7652 0.00883 0.00322 -0.0178 0.3136 1.0000
5.750 0.7889 0.00917 0.00345 -0.0171 0.2834 1.0000
6.000 0.8124 0.00953 0.00373 -0.0165 0.2546 1.0000
6.250 0.8353 0.00994 0.00403 -0.0157 0.2248 1.0000
6.750 0.8799 0.01091 0.00475 -0.0142 0.1654 1.0000
7.000 0.9017 0.01143 0.00517 -0.0133 0.1380 1.0000
7.250 0.9229 0.01199 0.00561 -0.0124 0.1114 1.0000
7.750 0.9635 0.01328 0.00666 -0.0103 0.0611 1.0000
8.000 0.9800 0.01435 0.00752 -0.0088 0.0300 1.0000
8.250 0.9980 0.01522 0.00840 -0.0072 0.0224 1.0000
8.500 1.0180 0.01582 0.00909 -0.0060 0.0207 1.0000
8.750 1.0369 0.01650 0.00982 -0.0047 0.0189 1.0000
9.000 1.0528 0.01744 0.01083 -0.0030 0.0172 1.0000
9.250 1.0658 0.01859 0.01210 -0.0010 0.0163 1.0000
9.500 1.0817 0.01941 0.01302 0.0007 0.0159 1.0000
9.750 1.0964 0.02030 0.01402 0.0024 0.0154 1.0000
10.000 1.1095 0.02129 0.01511 0.0043 0.0149 1.0000
10.250 1.1208 0.02238 0.01631 0.0064 0.0145 1.0000
10.500 1.1307 0.02353 0.01755 0.0086 0.0141 1.0000
10.750 1.1386 0.02467 0.01878 0.0110 0.0137 1.0000
11.000 1.1437 0.02573 0.01991 0.0137 0.0133 1.0000
11.250 1.1465 0.02710 0.02136 0.0162 0.0129 1.0000
11.500 1.1469 0.02893 0.02327 0.0183 0.0124 1.0000
11.750 1.1465 0.03122 0.02569 0.0202 0.0121 1.0000
12.000 1.1428 0.03452 0.02917 0.0222 0.0118 1.0000
12.250 1.1468 0.03639 0.03120 0.0231 0.0117 1.0000
12.500 1.1497 0.03837 0.03333 0.0236 0.0116 1.0000
12.750 1.1497 0.04106 0.03618 0.0243 0.0116 1.0000
13.000 1.1492 0.04364 0.03893 0.0244 0.0114 1.0000
13.250 1.1457 0.04680 0.04226 0.0243 0.0113 1.0000
13.500 1.1407 0.05023 0.04587 0.0238 0.0112 1.0000
13.750 1.1339 0.05401 0.04982 0.0229 0.0112 1.0000
14.000 1.1259 0.05807 0.05406 0.0214 0.0109 1.0000
14.250 1.1164 0.06253 0.05868 0.0197 0.0109 1.0000
14.500 1.1050 0.06748 0.06379 0.0175 0.0109 1.0000
14.750 1.0903 0.07319 0.06968 0.0147 0.0109 1.0000
15.000 1.0713 0.08006 0.07673 0.0111 0.0110 1.0000
15.250 1.0591 0.08609 0.08290 0.0076 0.0108 1.0000
15.500 1.0430 0.09326 0.09022 0.0033 0.0107 1.0000
15.750 1.0183 0.10246 0.09958 -0.0019 0.0112 1.0000
16.000 0.9983 0.11139 0.10866 -0.0073 0.0111 1.0000
16.250 0.9748 0.12145 0.11884 -0.0132 0.0114 1.0000
16.500 0.9424 0.13425 0.13177 -0.0205 0.0117 1.0000
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Polar data table (+)
Polar graphs
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