EPPLER 637 AIRFOIL (e637-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 637 AIRFOIL (e637-il) Reynolds number: 1,000,000 Max Cl/Cd: 126.77 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e637-il-1000000.txt Download as CSV file: xf-e637-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 637 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3273 0.08511 0.08269 -0.0285 0.7207 0.0102
-8.500 -0.3304 0.08109 0.07862 -0.0305 0.7013 0.0103
-8.250 -0.3346 0.07722 0.07471 -0.0327 0.6840 0.0103
-8.000 -0.3465 0.07291 0.07037 -0.0360 0.6695 0.0104
-7.750 -0.3601 0.06906 0.06646 -0.0364 0.6574 0.0104
-7.500 -0.3676 0.06429 0.06160 -0.0373 0.6462 0.0106
-7.250 -0.3674 0.06043 0.05763 -0.0374 0.6352 0.0106
-7.000 -0.3653 0.05616 0.05323 -0.0371 0.6245 0.0106
-6.750 -0.3607 0.05213 0.04905 -0.0363 0.6137 0.0106
-6.500 -0.3659 0.04616 0.04288 -0.0347 0.6062 0.0108
-6.000 -0.3447 0.04100 0.03743 -0.0324 0.5850 0.0110
-5.750 -0.3299 0.03942 0.03574 -0.0314 0.5747 0.0113
-5.500 -0.3151 0.03704 0.03319 -0.0301 0.5648 0.0115
-5.250 -0.2993 0.03457 0.03054 -0.0287 0.5557 0.0121
-5.000 -0.2919 0.02742 0.02278 -0.0239 0.5502 0.0139
-4.750 -0.2724 0.02623 0.02148 -0.0229 0.5404 0.0142
-4.500 -0.2522 0.02501 0.02011 -0.0219 0.5307 0.0145
-4.250 -0.2311 0.02369 0.01863 -0.0207 0.5217 0.0150
-4.000 -0.2073 0.02247 0.01695 -0.0186 0.5137 0.0169
-3.750 -0.1905 0.01920 0.01340 -0.0167 0.5065 0.0176
-3.500 -0.1666 0.01833 0.01245 -0.0161 0.4981 0.0181
-3.250 -0.1428 0.01463 0.00813 -0.0137 0.4916 0.0142
-3.000 -0.1166 0.01317 0.00651 -0.0129 0.4834 0.0124
-2.750 -0.0908 0.01214 0.00533 -0.0121 0.4760 0.0122
-2.500 -0.0654 0.01154 0.00461 -0.0113 0.4684 0.0124
-2.250 -0.0397 0.01110 0.00411 -0.0107 0.4620 0.0129
-2.000 -0.0140 0.01079 0.00372 -0.0100 0.4549 0.0133
-1.750 0.0115 0.01045 0.00333 -0.0094 0.4484 0.0137
-1.500 0.0358 0.01003 0.00286 -0.0085 0.4420 0.0143
-1.250 0.0614 0.00982 0.00261 -0.0080 0.4358 0.0150
-1.000 0.0878 0.00965 0.00242 -0.0075 0.4304 0.0160
-0.750 0.1142 0.00954 0.00226 -0.0071 0.4245 0.0170
-0.500 0.1399 0.00933 0.00202 -0.0065 0.4192 0.0187
-0.250 0.1665 0.00920 0.00187 -0.0061 0.4140 0.0217
0.000 0.1917 0.00899 0.00174 -0.0054 0.4088 0.0526
0.250 0.2144 0.00853 0.00168 -0.0045 0.4045 0.1917
0.500 0.2085 0.00647 0.00154 0.0021 0.4017 0.7821
0.750 0.2471 0.00630 0.00180 0.0004 0.3966 0.9351
1.000 0.2839 0.00665 0.00208 -0.0010 0.3907 0.9581
1.250 0.3511 0.00711 0.00246 -0.0090 0.3859 0.9659
1.500 0.3852 0.00736 0.00265 -0.0101 0.3815 0.9730
1.750 0.4351 0.00757 0.00276 -0.0147 0.3760 0.9749
2.000 0.4694 0.00761 0.00277 -0.0161 0.3727 0.9760
2.250 0.5023 0.00765 0.00277 -0.0172 0.3687 0.9772
2.500 0.5344 0.00772 0.00280 -0.0182 0.3649 0.9787
2.750 0.5649 0.00783 0.00285 -0.0188 0.3603 0.9806
3.000 0.5936 0.00788 0.00289 -0.0190 0.3577 0.9829
3.250 0.6223 0.00794 0.00294 -0.0193 0.3543 0.9849
3.500 0.6559 0.00798 0.00295 -0.0206 0.3508 0.9857
3.750 0.6887 0.00806 0.00298 -0.0218 0.3468 0.9866
4.000 0.7215 0.00811 0.00302 -0.0230 0.3436 0.9876
4.250 0.7539 0.00814 0.00306 -0.0241 0.3404 0.9888
4.500 0.7854 0.00820 0.00310 -0.0250 0.3371 0.9901
4.750 0.8162 0.00829 0.00317 -0.0258 0.3335 0.9917
5.000 0.8461 0.00841 0.00327 -0.0264 0.3297 0.9933
5.250 0.8763 0.00845 0.00334 -0.0270 0.3270 0.9946
5.500 0.9089 0.00848 0.00338 -0.0282 0.3237 0.9954
5.750 0.9411 0.00854 0.00344 -0.0293 0.3198 0.9964
6.000 0.9724 0.00867 0.00353 -0.0303 0.3153 0.9976
6.250 1.0042 0.00869 0.00359 -0.0313 0.3122 0.9987
6.500 1.0355 0.00874 0.00366 -0.0323 0.3077 0.9996
6.750 1.0625 0.00886 0.00377 -0.0324 0.3030 1.0000
7.000 1.0861 0.00897 0.00389 -0.0317 0.2990 1.0000
7.250 1.1101 0.00904 0.00400 -0.0311 0.2951 1.0000
7.500 1.1335 0.00916 0.00413 -0.0304 0.2911 1.0000
7.750 1.1562 0.00933 0.00429 -0.0296 0.2861 1.0000
8.000 1.1801 0.00941 0.00443 -0.0289 0.2825 1.0000
8.250 1.2032 0.00954 0.00457 -0.0282 0.2772 1.0000
8.500 1.2254 0.00973 0.00476 -0.0274 0.2714 1.0000
8.750 1.2487 0.00985 0.00492 -0.0266 0.2659 1.0000
9.000 1.2706 0.01006 0.00513 -0.0257 0.2597 1.0000
9.250 1.2930 0.01023 0.00533 -0.0249 0.2538 1.0000
9.500 1.3143 0.01047 0.00557 -0.0239 0.2456 1.0000
9.750 1.3353 0.01073 0.00582 -0.0229 0.2343 1.0000
10.000 1.3560 0.01100 0.00609 -0.0219 0.2261 1.0000
10.250 1.3751 0.01138 0.00643 -0.0206 0.2127 1.0000
10.500 1.3929 0.01182 0.00682 -0.0191 0.1971 1.0000
10.750 1.4091 0.01234 0.00728 -0.0174 0.1805 1.0000
11.000 1.4224 0.01302 0.00786 -0.0153 0.1605 1.0000
11.250 1.4342 0.01374 0.00849 -0.0130 0.1413 1.0000
11.500 1.4447 0.01449 0.00916 -0.0104 0.1257 1.0000
11.750 1.4539 0.01524 0.00986 -0.0077 0.1115 1.0000
12.000 1.4615 0.01599 0.01057 -0.0048 0.0991 1.0000
12.500 1.4697 0.01755 0.01206 0.0021 0.0793 1.0000
12.750 1.4658 0.01826 0.01279 0.0070 0.0724 1.0000
13.000 1.4614 0.01927 0.01379 0.0112 0.0653 1.0000
13.250 1.4587 0.02054 0.01507 0.0142 0.0586 1.0000
13.500 1.4584 0.02203 0.01659 0.0162 0.0535 1.0000
14.000 1.4539 0.02635 0.02099 0.0181 0.0434 1.0000
14.250 1.4480 0.02929 0.02397 0.0183 0.0386 1.0000
14.500 1.4432 0.03230 0.02703 0.0182 0.0347 1.0000
14.750 1.4339 0.03589 0.03067 0.0179 0.0306 1.0000
15.000 1.4244 0.03955 0.03439 0.0175 0.0276 1.0000
15.250 1.4095 0.04389 0.03879 0.0169 0.0236 1.0000
15.500 1.3958 0.04812 0.04308 0.0163 0.0213 1.0000
15.750 1.3822 0.05241 0.04744 0.0156 0.0188 1.0000
16.000 1.3659 0.05712 0.05220 0.0147 0.0167 1.0000
16.250 1.3556 0.06128 0.05644 0.0138 0.0154 1.0000
16.500 1.3407 0.06612 0.06134 0.0126 0.0138 1.0000
16.750 1.3297 0.07055 0.06585 0.0115 0.0128 1.0000
17.000 1.3178 0.07521 0.07057 0.0102 0.0117 1.0000
17.250 1.3055 0.07996 0.07539 0.0089 0.0110 1.0000
17.500 1.2982 0.08410 0.07960 0.0077 0.0109 1.0000
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