EPPLER 1233 AIRFOIL (e1233-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 1233 AIRFOIL (e1233-il) Reynolds number: 50,000 Max Cl/Cd: 15.75 at α=-1.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e1233-il-50000.txt Download as CSV file: xf-e1233-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 1233 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3769 0.12013 0.11535 -0.0050 1.0000 0.3432
-7.500 -0.3914 0.08421 0.07954 -0.0494 0.9583 0.1555
-7.250 -0.5687 0.08370 0.07852 -0.0385 0.9843 0.1406
-7.000 -0.5921 0.07402 0.06810 -0.0443 0.9726 0.1280
-6.750 -0.5735 0.06888 0.06257 -0.0479 0.9608 0.1263
-6.500 -0.5553 0.06415 0.05739 -0.0508 0.9488 0.1261
-6.250 -0.5361 0.05978 0.05237 -0.0532 0.9367 0.1267
-6.000 -0.5112 0.05597 0.04797 -0.0552 0.9245 0.1276
-5.750 -0.4799 0.05350 0.04560 -0.0571 0.9127 0.1334
-5.500 -0.4443 0.05067 0.04212 -0.0598 0.9018 0.1406
-5.250 -0.4109 0.04837 0.03970 -0.0615 0.8898 0.1504
-5.000 -0.3840 0.04656 0.03754 -0.0621 0.8774 0.1622
-4.750 -0.3469 0.04487 0.03578 -0.0639 0.8668 0.1810
-4.500 -0.3123 0.04346 0.03434 -0.0651 0.8552 0.2058
-4.250 -0.2885 0.04250 0.03336 -0.0646 0.8431 0.2311
-4.000 -0.2414 0.04090 0.03210 -0.0670 0.8345 0.2761
-3.750 -0.2292 0.04042 0.03182 -0.0648 0.8213 0.3093
-3.500 -0.2022 0.03916 0.03114 -0.0644 0.8114 0.3808
-3.250 -0.1933 0.03864 0.03212 -0.0592 0.8011 0.5210
-3.000 -0.1965 0.04106 0.03495 -0.0492 0.7902 0.6524
-2.750 -0.1863 0.04343 0.03717 -0.0405 0.7805 0.7244
-2.500 -0.1904 0.04504 0.03867 -0.0330 0.7694 0.7615
-2.250 -0.1657 0.04636 0.03974 -0.0275 0.7611 0.8080
-2.000 0.4796 0.03932 0.03043 -0.1144 0.7594 0.9721
-1.750 0.5391 0.03838 0.02928 -0.1222 0.7451 0.9926
-1.500 0.5775 0.03794 0.02867 -0.1258 0.7336 1.0000
-1.250 0.5974 0.03794 0.02853 -0.1256 0.7237 1.0000
-1.000 0.5833 0.03946 0.03007 -0.1202 0.7120 1.0000
-0.750 0.6107 0.03923 0.02968 -0.1210 0.7039 1.0000
-0.500 0.5773 0.04164 0.03218 -0.1129 0.6922 1.0000
-0.250 0.6072 0.04134 0.03171 -0.1139 0.6848 1.0000
0.000 0.5469 0.04496 0.03550 -0.1026 0.6736 1.0000
0.250 0.5625 0.04532 0.03575 -0.1015 0.6665 1.0000
0.500 0.4263 0.05179 0.04253 -0.0816 0.6552 1.0000
0.750 0.4165 0.05323 0.04393 -0.0779 0.6482 1.0000
1.000 0.4758 0.05186 0.04234 -0.0816 0.6435 1.0000
1.250 0.3209 0.05944 0.05026 -0.0637 0.6368 1.0000
1.500 0.2958 0.06140 0.05221 -0.0592 0.6326 1.0000
1.750 0.3006 0.06228 0.05300 -0.0572 0.6281 1.0000
2.000 0.3025 0.06337 0.05402 -0.0550 0.6241 1.0000
2.250 0.2656 0.06565 0.05632 -0.0497 0.6233 1.0000
2.500 0.2463 0.06742 0.05808 -0.0461 0.6240 1.0000
2.750 0.2362 0.06899 0.05961 -0.0433 0.6250 1.0000
3.250 0.0556 0.07721 0.06830 -0.0317 0.7580 1.0000
3.500 0.0541 0.07706 0.06809 -0.0298 0.7462 1.0000
3.750 0.0942 0.08069 0.07156 -0.0336 0.7400 1.0000
4.000 0.0868 0.08020 0.07102 -0.0312 0.7284 1.0000
4.250 0.1259 0.08367 0.07435 -0.0348 0.7215 1.0000
4.500 0.1214 0.08382 0.07445 -0.0331 0.7117 1.0000
4.750 0.1528 0.08656 0.07708 -0.0357 0.7039 1.0000
5.000 0.1590 0.08788 0.07833 -0.0355 0.6956 1.0000
5.250 0.1862 0.09022 0.08059 -0.0374 0.6856 1.0000
5.500 0.2045 0.09280 0.08309 -0.0387 0.6802 1.0000
5.750 0.2140 0.09384 0.08408 -0.0387 0.6690 1.0000
6.000 0.2538 0.09807 0.08821 -0.0421 0.6639 1.0000
6.250 0.2449 0.09792 0.08805 -0.0403 0.6526 1.0000
6.500 0.2803 0.10150 0.09154 -0.0430 0.6458 1.0000
6.750 0.2778 0.10237 0.09239 -0.0422 0.6370 1.0000
7.000 0.3010 0.10489 0.09485 -0.0436 0.6292 1.0000
7.500 0.3260 0.10897 0.09887 -0.0448 0.6132 1.0000
7.750 0.3614 0.11300 0.10284 -0.0473 0.6074 1.0000
8.000 0.3544 0.11359 0.10343 -0.0464 0.5987 1.0000
8.250 0.3763 0.11623 0.10604 -0.0476 0.5913 1.0000
8.500 0.4158 0.12152 0.11129 -0.0506 0.5875 1.0000
8.750 0.3977 0.12046 0.11025 -0.0487 0.5765 1.0000
9.000 0.4283 0.12418 0.11394 -0.0506 0.5703 1.0000
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Polar data table (+)
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