EPPLER 637 AIRFOIL (e637-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 637 AIRFOIL (e637-il) Reynolds number: 500,000 Max Cl/Cd: 97.98 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e637-il-500000.txt Download as CSV file: xf-e637-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3587 0.09419 0.09212 -0.0150 1.0000 0.0153
-8.750 -0.3570 0.09062 0.08858 -0.0167 1.0000 0.0156
-8.500 -0.3564 0.08689 0.08490 -0.0186 1.0000 0.0158
-8.250 -0.3450 0.08168 0.07969 -0.0250 0.9337 0.0162
-8.000 -0.3346 0.07437 0.07196 -0.0375 0.8187 0.0167
-7.750 -0.3483 0.07086 0.06824 -0.0378 0.7743 0.0167
-7.500 -0.3589 0.06608 0.06325 -0.0389 0.7502 0.0169
-7.250 -0.3600 0.06210 0.05908 -0.0388 0.7287 0.0170
-7.000 -0.3626 0.05787 0.05471 -0.0380 0.7109 0.0171
-6.750 -0.3573 0.05534 0.05206 -0.0371 0.6932 0.0173
-6.500 -0.3491 0.05312 0.04971 -0.0363 0.6769 0.0176
-6.250 -0.3395 0.05052 0.04698 -0.0355 0.6622 0.0178
-6.000 -0.3285 0.04793 0.04423 -0.0346 0.6486 0.0182
-5.750 -0.3164 0.04513 0.04125 -0.0335 0.6361 0.0192
-5.500 -0.3079 0.03899 0.03449 -0.0301 0.6281 0.0212
-5.250 -0.2933 0.03699 0.03245 -0.0294 0.6151 0.0216
-5.000 -0.2766 0.03542 0.03076 -0.0285 0.6030 0.0221
-4.750 -0.2589 0.03363 0.02879 -0.0273 0.5923 0.0230
-4.500 -0.2433 0.03029 0.02475 -0.0235 0.5836 0.0260
-4.250 -0.2239 0.02824 0.02270 -0.0230 0.5733 0.0267
-4.000 -0.2030 0.02719 0.02156 -0.0223 0.5629 0.0278
-3.750 -0.1778 0.02754 0.02141 -0.0201 0.5526 0.0315
-3.500 -0.1611 0.02366 0.01742 -0.0191 0.5443 0.0329
-3.250 -0.1381 0.02249 0.01614 -0.0183 0.5347 0.0342
-3.000 -0.1112 0.02315 0.01648 -0.0170 0.5258 0.0385
-2.750 -0.0860 0.01624 0.00879 -0.0140 0.5199 0.0203
-2.500 -0.0592 0.01483 0.00720 -0.0134 0.5120 0.0195
-2.250 -0.0327 0.01393 0.00616 -0.0127 0.5036 0.0196
-2.000 -0.0067 0.01328 0.00539 -0.0120 0.4959 0.0200
-1.750 0.0189 0.01279 0.00481 -0.0113 0.4883 0.0205
-1.500 0.0431 0.01226 0.00423 -0.0104 0.4817 0.0218
-1.250 0.0680 0.01190 0.00386 -0.0097 0.4747 0.0232
-1.000 0.0926 0.01163 0.00350 -0.0088 0.4683 0.0243
-0.750 0.1181 0.01134 0.00318 -0.0081 0.4617 0.0259
-0.500 0.1429 0.01107 0.00288 -0.0074 0.4556 0.0292
-0.250 0.1682 0.01086 0.00265 -0.0067 0.4502 0.0408
0.000 0.1739 0.00895 0.00242 -0.0029 0.4458 0.5314
0.250 0.2278 0.00829 0.00299 -0.0072 0.4389 0.9438
0.500 0.3201 0.00929 0.00381 -0.0194 0.4302 0.9718
0.750 0.3882 0.00958 0.00392 -0.0276 0.4234 0.9808
1.000 0.4383 0.00962 0.00387 -0.0322 0.4175 0.9873
1.250 0.4732 0.00958 0.00377 -0.0337 0.4124 0.9897
1.500 0.5060 0.00962 0.00370 -0.0349 0.4077 0.9924
1.750 0.5384 0.00964 0.00367 -0.0360 0.4035 0.9947
2.000 0.5728 0.00957 0.00357 -0.0375 0.3992 0.9965
2.250 0.6061 0.00955 0.00349 -0.0389 0.3946 0.9985
2.500 0.6377 0.00961 0.00346 -0.0398 0.3903 1.0000
2.750 0.6616 0.00963 0.00350 -0.0392 0.3868 1.0000
3.000 0.6855 0.00968 0.00354 -0.0385 0.3831 1.0000
3.250 0.7092 0.00975 0.00358 -0.0378 0.3794 1.0000
3.500 0.7326 0.00991 0.00366 -0.0371 0.3756 1.0000
3.750 0.7566 0.00995 0.00373 -0.0364 0.3723 1.0000
4.000 0.7806 0.01001 0.00381 -0.0358 0.3689 1.0000
4.250 0.8044 0.01009 0.00389 -0.0351 0.3655 1.0000
4.500 0.8278 0.01024 0.00398 -0.0344 0.3619 1.0000
4.750 0.8516 0.01036 0.00411 -0.0337 0.3587 1.0000
5.000 0.8755 0.01043 0.00423 -0.0330 0.3553 1.0000
5.250 0.8992 0.01052 0.00433 -0.0323 0.3519 1.0000
5.500 0.9227 0.01065 0.00444 -0.0316 0.3486 1.0000
5.750 0.9459 0.01086 0.00463 -0.0309 0.3449 1.0000
6.000 0.9697 0.01093 0.00477 -0.0302 0.3418 1.0000
6.250 0.9932 0.01103 0.00491 -0.0295 0.3382 1.0000
6.500 1.0164 0.01115 0.00505 -0.0287 0.3346 1.0000
6.750 1.0391 0.01139 0.00525 -0.0279 0.3307 1.0000
7.000 1.0623 0.01148 0.00543 -0.0271 0.3270 1.0000
7.250 1.0853 0.01158 0.00558 -0.0263 0.3226 1.0000
7.500 1.1079 0.01174 0.00575 -0.0255 0.3184 1.0000
7.750 1.1301 0.01195 0.00597 -0.0246 0.3142 1.0000
8.000 1.1529 0.01205 0.00616 -0.0238 0.3098 1.0000
8.250 1.1752 0.01220 0.00635 -0.0229 0.3054 1.0000
8.500 1.1965 0.01247 0.00660 -0.0219 0.3008 1.0000
8.750 1.2191 0.01257 0.00682 -0.0211 0.2962 1.0000
9.000 1.2409 0.01272 0.00703 -0.0201 0.2909 1.0000
9.250 1.2616 0.01296 0.00728 -0.0190 0.2854 1.0000
9.500 1.2836 0.01310 0.00751 -0.0181 0.2794 1.0000
9.750 1.3038 0.01335 0.00778 -0.0170 0.2734 1.0000
10.000 1.3249 0.01355 0.00807 -0.0159 0.2674 1.0000
10.250 1.3444 0.01381 0.00835 -0.0147 0.2599 1.0000
10.500 1.3643 0.01405 0.00866 -0.0135 0.2509 1.0000
10.750 1.3823 0.01439 0.00901 -0.0121 0.2413 1.0000
11.000 1.3989 0.01479 0.00941 -0.0105 0.2298 1.0000
11.250 1.4143 0.01525 0.00987 -0.0087 0.2162 1.0000
11.500 1.4279 0.01578 0.01040 -0.0067 0.2010 1.0000
11.750 1.4385 0.01645 0.01103 -0.0043 0.1840 1.0000
12.000 1.4448 0.01728 0.01180 -0.0013 0.1662 1.0000
12.250 1.4473 0.01822 0.01269 0.0022 0.1506 1.0000
12.500 1.4433 0.01912 0.01356 0.0068 0.1384 1.0000
12.750 1.4387 0.02018 0.01461 0.0110 0.1272 1.0000
13.000 1.4341 0.02154 0.01598 0.0143 0.1168 1.0000
13.250 1.4292 0.02331 0.01776 0.0167 0.1076 1.0000
13.500 1.4232 0.02558 0.02005 0.0181 0.0995 1.0000
13.750 1.4187 0.02810 0.02262 0.0188 0.0915 1.0000
14.000 1.4129 0.03099 0.02556 0.0190 0.0851 1.0000
14.250 1.4032 0.03447 0.02908 0.0189 0.0784 1.0000
14.500 1.3957 0.03782 0.03250 0.0187 0.0731 1.0000
14.750 1.3829 0.04182 0.03655 0.0183 0.0679 1.0000
15.000 1.3725 0.04562 0.04042 0.0178 0.0640 1.0000
15.250 1.3598 0.04973 0.04459 0.0172 0.0597 1.0000
15.500 1.3441 0.05424 0.04915 0.0164 0.0558 1.0000
15.750 1.3340 0.05820 0.05318 0.0157 0.0526 1.0000
16.000 1.3197 0.06285 0.05788 0.0147 0.0488 1.0000
16.250 1.3075 0.06732 0.06240 0.0136 0.0454 1.0000
16.500 1.2960 0.07183 0.06697 0.0124 0.0419 1.0000
16.750 1.2827 0.07663 0.07181 0.0111 0.0384 1.0000
17.000 1.2698 0.08148 0.07670 0.0097 0.0341 1.0000
17.250 1.2565 0.08651 0.08177 0.0082 0.0301 1.0000
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