EPPLER 548 AIRFOIL (e548-il) Xfoil prediction polar at RE=200,000 Ncrit=9
Details | Polar file |
---|---|
Airfoil: EPPLER 548 AIRFOIL (e548-il) Reynolds number: 200,000 Max Cl/Cd: 61.76 at α=10.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e548-il-200000.txt Download as CSV file: xf-e548-il-200000.csv |
XFOIL Version 6.96 Calculated polar for: EPPLER 548 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.500 -0.3380 0.09391 0.09084 -0.0484 1.0000 0.0599 -11.250 -0.3688 0.08355 0.08050 -0.0549 1.0000 0.0609 -11.000 -0.4010 0.07530 0.07221 -0.0591 1.0000 0.0608 -10.750 -0.4346 0.06867 0.06551 -0.0612 1.0000 0.0607 -10.500 -0.6219 0.07028 0.06644 -0.0569 1.0000 0.0540 -10.250 -0.6700 0.07275 0.06868 -0.0489 1.0000 0.0534 -9.500 -0.6578 0.03572 0.03070 -0.0457 0.9719 0.0276 -9.250 -0.6235 0.03105 0.02602 -0.0488 0.9571 0.0258 -9.000 -0.6816 0.03743 0.03128 -0.0400 0.9605 0.0227 -8.750 -0.6435 0.03351 0.02702 -0.0433 0.9447 0.0223 -8.500 -0.5971 0.03006 0.02313 -0.0475 0.9279 0.0222 -8.250 -0.5443 0.02744 0.01998 -0.0526 0.9064 0.0229 -8.000 -0.4875 0.02528 0.01779 -0.0584 0.8807 0.0252 -7.750 -0.4362 0.02368 0.01585 -0.0623 0.8528 0.0272 -7.500 -0.3921 0.02225 0.01439 -0.0645 0.8285 0.0295 -7.250 -0.3676 0.02152 0.01352 -0.0638 0.8048 0.0314 -7.000 -0.3505 0.02068 0.01260 -0.0621 0.7850 0.0339 -6.750 -0.3360 0.02007 0.01189 -0.0602 0.7681 0.0389 -6.500 -0.3252 0.01933 0.01109 -0.0577 0.7535 0.0451 -6.250 -0.3191 0.01854 0.01026 -0.0546 0.7408 0.0533 -6.000 -0.3186 0.01771 0.00952 -0.0505 0.7287 0.0695 -5.750 -0.3181 0.01667 0.00872 -0.0465 0.7177 0.1127 -5.500 -0.3206 0.01542 0.00795 -0.0424 0.7089 0.2058 -5.250 -0.3300 0.01372 0.00711 -0.0373 0.6992 0.3702 -5.000 -0.3241 0.01367 0.00819 -0.0323 0.6918 0.6137 -4.750 -0.2994 0.01463 0.00905 -0.0306 0.6826 0.6579 -4.500 -0.2691 0.01607 0.01034 -0.0291 0.6756 0.6800 -4.250 -0.2385 0.01726 0.01145 -0.0278 0.6674 0.6932 -4.000 -0.2097 0.01817 0.01221 -0.0266 0.6607 0.7055 -3.750 -0.1752 0.01946 0.01342 -0.0257 0.6537 0.7148 -3.500 -0.1424 0.02040 0.01424 -0.0248 0.6469 0.7233 -3.250 -0.1114 0.02143 0.01512 -0.0237 0.6418 0.7368 -3.000 -0.0441 0.02364 0.01726 -0.0260 0.6348 0.7447 -2.750 -0.0309 0.02337 0.01690 -0.0239 0.6298 0.7542 -2.500 0.0058 0.02354 0.01689 -0.0250 0.6252 0.7568 -2.250 0.0328 0.02348 0.01679 -0.0248 0.6197 0.7608 -2.000 0.0324 0.02281 0.01607 -0.0214 0.6152 0.7696 -1.750 0.0653 0.02279 0.01594 -0.0220 0.6108 0.7713 -1.500 0.0950 0.02277 0.01582 -0.0223 0.6071 0.7737 -1.250 0.1172 0.02259 0.01563 -0.0217 0.6026 0.7770 -1.000 0.1262 0.02211 0.01510 -0.0197 0.5987 0.7825 -0.750 0.1440 0.02173 0.01463 -0.0188 0.5951 0.7860 -0.500 0.1730 0.02169 0.01448 -0.0190 0.5917 0.7876 -0.250 0.1967 0.02155 0.01437 -0.0186 0.5875 0.7895 0.000 0.2200 0.02141 0.01422 -0.0182 0.5838 0.7918 0.250 0.2417 0.02120 0.01396 -0.0178 0.5805 0.7941 0.500 0.2631 0.02096 0.01365 -0.0175 0.5775 0.7969 0.750 0.2845 0.02072 0.01332 -0.0176 0.5746 0.7995 1.000 0.3055 0.02048 0.01313 -0.0173 0.5706 0.8016 1.250 0.3303 0.02042 0.01309 -0.0170 0.5671 0.8029 1.500 0.3558 0.02034 0.01300 -0.0169 0.5640 0.8041 1.750 0.3823 0.02026 0.01287 -0.0172 0.5611 0.8053 2.000 0.4094 0.02029 0.01283 -0.0175 0.5584 0.8068 2.250 0.4319 0.02027 0.01291 -0.0171 0.5547 0.8088 2.500 0.4564 0.02025 0.01293 -0.0172 0.5513 0.8104 2.750 0.4826 0.02019 0.01286 -0.0177 0.5479 0.8119 3.000 0.5104 0.02014 0.01278 -0.0183 0.5450 0.8134 3.250 0.5397 0.02019 0.01278 -0.0194 0.5425 0.8151 3.500 0.5642 0.02027 0.01294 -0.0197 0.5387 0.8172 3.750 0.5881 0.02030 0.01305 -0.0195 0.5349 0.8183 4.000 0.6142 0.02033 0.01312 -0.0196 0.5314 0.8195 4.250 0.6420 0.02035 0.01313 -0.0199 0.5284 0.8206 4.500 0.6710 0.02047 0.01322 -0.0205 0.5256 0.8219 4.750 0.6924 0.02062 0.01354 -0.0200 0.5210 0.8234 5.000 0.7177 0.02069 0.01368 -0.0200 0.5167 0.8252 5.250 0.7461 0.02070 0.01369 -0.0205 0.5130 0.8272 5.500 0.7772 0.02075 0.01370 -0.0216 0.5098 0.8290 5.750 0.7995 0.02095 0.01406 -0.0214 0.5048 0.8308 6.000 0.8263 0.02102 0.01420 -0.0218 0.5000 0.8325 6.250 0.8556 0.02095 0.01414 -0.0224 0.4960 0.8339 6.500 0.8832 0.02100 0.01421 -0.0226 0.4920 0.8354 6.750 0.9031 0.02112 0.01451 -0.0217 0.4859 0.8371 7.000 0.9307 0.02103 0.01446 -0.0219 0.4810 0.8389 7.250 0.9629 0.02096 0.01433 -0.0228 0.4768 0.8410 7.500 0.9805 0.02111 0.01471 -0.0216 0.4699 0.8437 7.750 1.0092 0.02095 0.01459 -0.0220 0.4643 0.8461 8.000 1.0379 0.02091 0.01457 -0.0226 0.4586 0.8482 8.250 1.0586 0.02086 0.01469 -0.0217 0.4512 0.8503 8.500 1.0891 0.02057 0.01436 -0.0221 0.4454 0.8522 8.750 1.1048 0.02060 0.01462 -0.0204 0.4372 0.8548 9.000 1.1329 0.02031 0.01432 -0.0205 0.4302 0.8573 9.250 1.1499 0.02032 0.01452 -0.0191 0.4211 0.8604 9.500 1.1765 0.02006 0.01426 -0.0190 0.4127 0.8636 9.750 1.1898 0.02001 0.01441 -0.0169 0.4020 0.8670 10.000 1.2044 0.01993 0.01446 -0.0148 0.3911 0.8707 10.250 1.2187 0.01987 0.01448 -0.0127 0.3786 0.8748 10.500 1.2308 0.01993 0.01462 -0.0106 0.3643 0.8792 10.750 1.2349 0.02005 0.01480 -0.0069 0.3495 0.8839 11.000 1.2367 0.02041 0.01522 -0.0031 0.3326 0.8898 11.250 1.2376 0.02103 0.01584 0.0002 0.3139 0.8966 11.500 1.2328 0.02189 0.01669 0.0041 0.2954 0.9046 11.750 1.2261 0.02311 0.01788 0.0075 0.2772 0.9152 12.000 1.2184 0.02462 0.01941 0.0105 0.2590 0.9303 12.250 1.2158 0.02667 0.02150 0.0111 0.2390 0.9594 12.500 1.2114 0.02923 0.02400 0.0110 0.2205 1.0000 12.750 1.2063 0.03205 0.02675 0.0109 0.2042 1.0000 13.000 1.1998 0.03505 0.02970 0.0108 0.1889 1.0000 13.250 1.1938 0.03808 0.03268 0.0107 0.1749 1.0000 13.500 1.1874 0.04118 0.03574 0.0106 0.1618 1.0000 13.750 1.1813 0.04430 0.03885 0.0105 0.1497 1.0000 14.000 1.1751 0.04752 0.04203 0.0102 0.1385 1.0000 14.250 1.1684 0.05087 0.04534 0.0099 0.1282 1.0000 14.500 1.1626 0.05424 0.04867 0.0095 0.1184 1.0000 14.750 1.1593 0.05746 0.05191 0.0090 0.1086 1.0000 15.000 1.1549 0.06085 0.05528 0.0084 0.1002 1.0000 15.250 1.1493 0.06445 0.05881 0.0076 0.0924 1.0000 15.500 1.1476 0.06774 0.06216 0.0069 0.0845 1.0000 15.750 1.1440 0.07122 0.06559 0.0061 0.0781 1.0000 16.000 1.1417 0.07474 0.06915 0.0051 0.0718 1.0000 16.250 1.1400 0.07812 0.07255 0.0043 0.0663 1.0000 16.500 1.1380 0.08171 0.07616 0.0031 0.0613 1.0000 16.750 1.1377 0.08500 0.07946 0.0022 0.0569 1.0000 17.000 1.1365 0.08861 0.08313 0.0010 0.0528 1.0000 17.250 1.1366 0.09185 0.08635 0.0000 0.0492 1.0000 |
Polar data table (+)
Polar graphs
<< Back to EPPLER 548 AIRFOIL (e548-il)