E226 (10.19%) (e226-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: E226 (10.19%) (e226-il) Reynolds number: 500,000 Max Cl/Cd: 77.08 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e226-il-500000.txt Download as CSV file: xf-e226-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: E226 (10.19%)
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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.5827 0.10596 0.10363 -0.0094 1.0000 0.0256
-11.000 -0.5833 0.10138 0.09907 -0.0116 1.0000 0.0262
-8.500 -0.7628 0.03048 0.02578 -0.0325 1.0000 0.0177
-8.250 -0.7511 0.02693 0.02208 -0.0308 1.0000 0.0167
-8.000 -0.7379 0.02479 0.01968 -0.0287 1.0000 0.0166
-7.750 -0.7245 0.02240 0.01702 -0.0265 1.0000 0.0161
-7.500 -0.7096 0.02044 0.01484 -0.0244 1.0000 0.0157
-7.250 -0.6940 0.01888 0.01310 -0.0222 1.0000 0.0154
-7.000 -0.6784 0.01767 0.01176 -0.0200 1.0000 0.0154
-6.750 -0.6630 0.01665 0.01065 -0.0178 1.0000 0.0156
-6.500 -0.6476 0.01579 0.00970 -0.0156 1.0000 0.0159
-6.250 -0.6211 0.01490 0.00872 -0.0157 0.9984 0.0161
-6.000 -0.5861 0.01402 0.00776 -0.0175 0.9954 0.0166
-5.750 -0.5517 0.01328 0.00695 -0.0191 0.9920 0.0171
-5.500 -0.5168 0.01257 0.00616 -0.0208 0.9882 0.0177
-5.250 -0.4811 0.01175 0.00528 -0.0228 0.9852 0.0195
-5.000 -0.4464 0.01127 0.00475 -0.0243 0.9811 0.0212
-4.750 -0.4108 0.01082 0.00423 -0.0260 0.9768 0.0238
-4.500 -0.3752 0.00991 0.00360 -0.0279 0.9735 0.0728
-4.250 -0.3431 0.00919 0.00323 -0.0292 0.9682 0.1458
-4.000 -0.3116 0.00865 0.00297 -0.0303 0.9616 0.2100
-3.750 -0.2802 0.00824 0.00275 -0.0312 0.9552 0.2641
-3.500 -0.2521 0.00788 0.00257 -0.0313 0.9461 0.3163
-3.250 -0.2250 0.00758 0.00241 -0.0311 0.9364 0.3624
-3.000 -0.1986 0.00727 0.00226 -0.0307 0.9269 0.4169
-2.750 -0.1742 0.00698 0.00216 -0.0299 0.9154 0.4741
-2.500 -0.1491 0.00678 0.00206 -0.0292 0.9037 0.5161
-2.250 -0.1236 0.00662 0.00196 -0.0285 0.8922 0.5524
-2.000 -0.0982 0.00647 0.00188 -0.0278 0.8805 0.5880
-1.750 -0.0727 0.00633 0.00181 -0.0270 0.8685 0.6234
-1.500 -0.0471 0.00622 0.00176 -0.0263 0.8553 0.6581
-1.250 -0.0215 0.00611 0.00171 -0.0256 0.8421 0.6907
-1.000 0.0042 0.00602 0.00168 -0.0249 0.8287 0.7223
-0.750 0.0299 0.00595 0.00165 -0.0242 0.8150 0.7525
-0.500 0.0557 0.00590 0.00164 -0.0235 0.8009 0.7816
-0.250 0.0814 0.00586 0.00163 -0.0227 0.7865 0.8094
0.000 0.1070 0.00584 0.00163 -0.0220 0.7717 0.8356
0.250 0.1325 0.00583 0.00165 -0.0211 0.7564 0.8602
0.500 0.1582 0.00585 0.00166 -0.0203 0.7405 0.8839
0.750 0.1842 0.00588 0.00169 -0.0196 0.7241 0.9055
1.000 0.2112 0.00594 0.00172 -0.0191 0.7071 0.9258
1.250 0.2405 0.00603 0.00176 -0.0191 0.6892 0.9430
1.500 0.2724 0.00612 0.00180 -0.0197 0.6699 0.9572
1.750 0.3085 0.00624 0.00185 -0.0213 0.6491 0.9666
2.000 0.3447 0.00636 0.00191 -0.0230 0.6271 0.9749
2.250 0.3827 0.00650 0.00196 -0.0252 0.6033 0.9819
2.500 0.4218 0.00664 0.00202 -0.0276 0.5775 0.9873
2.750 0.4604 0.00679 0.00208 -0.0300 0.5502 0.9930
3.000 0.4998 0.00694 0.00214 -0.0326 0.5214 0.9980
3.250 0.5291 0.00710 0.00221 -0.0331 0.4936 1.0000
3.500 0.5506 0.00728 0.00230 -0.0319 0.4680 1.0000
3.750 0.5724 0.00748 0.00241 -0.0308 0.4401 1.0000
4.000 0.5941 0.00772 0.00254 -0.0297 0.4075 1.0000
4.250 0.6159 0.00799 0.00270 -0.0286 0.3733 1.0000
4.500 0.6375 0.00832 0.00287 -0.0275 0.3368 1.0000
4.750 0.6592 0.00868 0.00308 -0.0264 0.2971 1.0000
5.000 0.6808 0.00908 0.00331 -0.0254 0.2580 1.0000
5.250 0.7024 0.00950 0.00358 -0.0243 0.2211 1.0000
5.500 0.7242 0.00993 0.00386 -0.0233 0.1866 1.0000
5.750 0.7459 0.01041 0.00418 -0.0223 0.1522 1.0000
6.000 0.7674 0.01093 0.00454 -0.0214 0.1180 1.0000
6.250 0.7889 0.01150 0.00494 -0.0204 0.0861 1.0000
6.500 0.8094 0.01220 0.00544 -0.0193 0.0507 1.0000
6.750 0.8282 0.01317 0.00619 -0.0179 0.0220 1.0000
7.000 0.8505 0.01372 0.00677 -0.0170 0.0185 1.0000
7.250 0.8717 0.01444 0.00755 -0.0158 0.0165 1.0000
7.500 0.8936 0.01503 0.00823 -0.0149 0.0157 1.0000
7.750 0.9147 0.01570 0.00897 -0.0138 0.0148 1.0000
8.000 0.9353 0.01640 0.00973 -0.0127 0.0141 1.0000
8.250 0.9547 0.01721 0.01059 -0.0115 0.0134 1.0000
8.500 0.9699 0.01847 0.01193 -0.0097 0.0128 1.0000
8.750 0.9850 0.01976 0.01332 -0.0080 0.0125 1.0000
9.000 1.0028 0.02076 0.01442 -0.0066 0.0124 1.0000
9.250 1.0207 0.02173 0.01549 -0.0053 0.0122 1.0000
9.500 1.0372 0.02291 0.01678 -0.0038 0.0121 1.0000
9.750 1.0533 0.02417 0.01816 -0.0023 0.0118 1.0000
10.000 1.0683 0.02560 0.01972 -0.0007 0.0117 1.0000
10.250 1.0824 0.02718 0.02147 0.0009 0.0116 1.0000
10.500 1.0946 0.02894 0.02341 0.0027 0.0114 1.0000
10.750 1.1045 0.03098 0.02565 0.0046 0.0115 1.0000
11.000 1.1098 0.03314 0.02802 0.0070 0.0116 1.0000
11.250 1.1097 0.03549 0.03061 0.0100 0.0117 1.0000
11.500 1.1052 0.03814 0.03350 0.0128 0.0118 1.0000
11.750 1.0967 0.04114 0.03674 0.0154 0.0119 1.0000
12.000 1.0861 0.04439 0.04021 0.0172 0.0121 1.0000
12.250 1.0718 0.04835 0.04437 0.0183 0.0123 1.0000
12.500 1.0558 0.05249 0.04872 0.0185 0.0123 1.0000
12.750 1.0540 0.05627 0.05257 0.0185 0.0127 1.0000
13.000 1.0435 0.05991 0.05637 0.0175 0.0128 1.0000
13.250 1.0280 0.06473 0.06136 0.0154 0.0128 1.0000
13.500 0.9093 0.06795 0.06496 0.0111 0.0128 1.0000
13.750 0.8904 0.07494 0.07209 0.0075 0.0129 1.0000
14.000 0.8700 0.08265 0.07996 0.0031 0.0130 1.0000
14.250 0.8449 0.09170 0.08919 -0.0025 0.0131 1.0000
14.500 0.7132 0.12457 0.12257 -0.0241 0.0143 1.0000
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Polar data table (+)
Polar graphs
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