EPPLER 637 AIRFOIL (e637-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 637 AIRFOIL (e637-il) Reynolds number: 200,000 Max Cl/Cd: 67.19 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e637-il-200000-n5.txt Download as CSV file: xf-e637-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.2833 0.10738 0.10432 -0.0116 1.0000 0.0239
-10.000 -0.2859 0.10337 0.10034 -0.0139 1.0000 0.0243
-9.750 -0.2891 0.09919 0.09619 -0.0164 1.0000 0.0245
-9.500 -0.2854 0.09393 0.09090 -0.0209 0.9020 0.0246
-9.250 -0.2706 0.09012 0.08685 -0.0210 0.8214 0.0251
-9.000 -0.2664 0.08691 0.08351 -0.0213 0.7858 0.0255
-8.750 -0.2630 0.08388 0.08038 -0.0216 0.7588 0.0268
-8.500 -0.2638 0.08014 0.07656 -0.0227 0.7394 0.0274
-8.250 -0.3330 0.08313 0.07972 -0.0293 0.8298 0.0256
-8.000 -0.3324 0.07941 0.07585 -0.0320 0.7843 0.0258
-7.500 -0.3427 0.07279 0.06901 -0.0345 0.7324 0.0263
-7.250 -0.3448 0.06968 0.06578 -0.0350 0.7137 0.0270
-7.000 -0.3448 0.06621 0.06218 -0.0356 0.6975 0.0278
-6.750 -0.3425 0.06268 0.05848 -0.0359 0.6834 0.0284
-6.500 -0.3408 0.05827 0.05366 -0.0362 0.6722 0.0297
-6.250 -0.3350 0.05485 0.04990 -0.0349 0.6599 0.0298
-6.000 -0.3259 0.05119 0.04610 -0.0340 0.6476 0.0300
-5.750 -0.3146 0.04815 0.04294 -0.0331 0.6356 0.0303
-5.250 -0.2938 0.03715 0.03113 -0.0281 0.6179 0.0162
-5.000 -0.2780 0.03498 0.02872 -0.0267 0.6073 0.0160
-4.750 -0.2614 0.03246 0.02591 -0.0251 0.5968 0.0157
-4.500 -0.2433 0.03002 0.02311 -0.0233 0.5869 0.0154
-4.250 -0.2243 0.02738 0.02005 -0.0214 0.5775 0.0152
-4.000 -0.2032 0.02501 0.01725 -0.0197 0.5682 0.0151
-3.750 -0.1803 0.02302 0.01482 -0.0182 0.5597 0.0152
-3.500 -0.1556 0.02132 0.01273 -0.0171 0.5505 0.0155
-3.250 -0.1301 0.02005 0.01109 -0.0162 0.5419 0.0167
-3.000 -0.1047 0.01930 0.01024 -0.0157 0.5326 0.0175
-2.750 -0.0783 0.01835 0.00908 -0.0151 0.5242 0.0179
-2.500 -0.0521 0.01747 0.00804 -0.0145 0.5160 0.0183
-2.250 -0.0264 0.01672 0.00717 -0.0138 0.5087 0.0189
-2.000 -0.0012 0.01611 0.00644 -0.0130 0.5011 0.0196
-1.750 0.0231 0.01558 0.00583 -0.0121 0.4941 0.0204
-1.500 0.0474 0.01514 0.00536 -0.0113 0.4867 0.0216
-1.000 0.0968 0.01455 0.00466 -0.0098 0.4738 0.0272
-0.750 0.1215 0.01429 0.00430 -0.0090 0.4678 0.0309
-0.500 0.1461 0.01403 0.00401 -0.0082 0.4621 0.0423
-0.250 0.1687 0.01355 0.00383 -0.0072 0.4560 0.1210
0.250 0.3456 0.01240 0.00495 -0.0299 0.4388 0.9813
0.500 0.3821 0.01245 0.00481 -0.0318 0.4334 0.9865
0.750 0.4152 0.01243 0.00470 -0.0330 0.4278 0.9892
1.000 0.4470 0.01245 0.00461 -0.0340 0.4227 0.9921
1.250 0.4779 0.01250 0.00453 -0.0348 0.4185 0.9947
1.500 0.5103 0.01249 0.00445 -0.0359 0.4133 0.9970
1.750 0.5421 0.01251 0.00441 -0.0369 0.4085 0.9993
2.000 0.5682 0.01258 0.00439 -0.0367 0.4042 1.0000
2.250 0.5917 0.01268 0.00442 -0.0360 0.4007 1.0000
2.500 0.6154 0.01276 0.00449 -0.0353 0.3964 1.0000
2.750 0.6390 0.01285 0.00455 -0.0346 0.3922 1.0000
3.000 0.6625 0.01297 0.00461 -0.0339 0.3885 1.0000
3.250 0.6860 0.01309 0.00469 -0.0331 0.3850 1.0000
3.500 0.7097 0.01321 0.00482 -0.0325 0.3812 1.0000
3.750 0.7332 0.01333 0.00494 -0.0317 0.3774 1.0000
4.000 0.7565 0.01347 0.00505 -0.0310 0.3737 1.0000
4.250 0.7798 0.01364 0.00518 -0.0302 0.3706 1.0000
4.500 0.8032 0.01379 0.00538 -0.0295 0.3668 1.0000
4.750 0.8264 0.01395 0.00558 -0.0287 0.3632 1.0000
5.000 0.8494 0.01412 0.00575 -0.0279 0.3596 1.0000
5.250 0.8723 0.01432 0.00591 -0.0271 0.3565 1.0000
5.500 0.8953 0.01451 0.00616 -0.0263 0.3530 1.0000
5.750 0.9182 0.01471 0.00643 -0.0255 0.3494 1.0000
6.000 0.9409 0.01490 0.00666 -0.0247 0.3457 1.0000
6.250 0.9634 0.01512 0.00687 -0.0239 0.3425 1.0000
6.500 0.9859 0.01536 0.00716 -0.0230 0.3392 1.0000
6.750 1.0083 0.01559 0.00748 -0.0221 0.3354 1.0000
7.000 1.0304 0.01581 0.00777 -0.0212 0.3315 1.0000
7.250 1.0524 0.01605 0.00803 -0.0203 0.3280 1.0000
7.500 1.0743 0.01632 0.00834 -0.0194 0.3245 1.0000
7.750 1.0957 0.01656 0.00871 -0.0184 0.3200 1.0000
8.000 1.1170 0.01680 0.00903 -0.0174 0.3155 1.0000
8.250 1.1381 0.01706 0.00929 -0.0164 0.3113 1.0000
8.500 1.1587 0.01732 0.00970 -0.0153 0.3060 1.0000
8.750 1.1791 0.01757 0.01004 -0.0142 0.3006 1.0000
9.000 1.1992 0.01786 0.01033 -0.0130 0.2961 1.0000
9.250 1.2189 0.01814 0.01080 -0.0118 0.2901 1.0000
9.500 1.2381 0.01843 0.01116 -0.0105 0.2846 1.0000
9.750 1.2569 0.01875 0.01157 -0.0092 0.2790 1.0000
10.000 1.2750 0.01907 0.01201 -0.0078 0.2721 1.0000
10.250 1.2922 0.01942 0.01243 -0.0063 0.2655 1.0000
10.500 1.3087 0.01979 0.01291 -0.0047 0.2574 1.0000
10.750 1.3242 0.02021 0.01341 -0.0030 0.2499 1.0000
11.000 1.3378 0.02067 0.01394 -0.0010 0.2411 1.0000
11.250 1.3510 0.02117 0.01456 0.0009 0.2321 1.0000
11.500 1.3610 0.02176 0.01520 0.0033 0.2226 1.0000
11.750 1.3682 0.02245 0.01593 0.0060 0.2129 1.0000
12.000 1.3709 0.02318 0.01673 0.0093 0.2024 1.0000
12.250 1.3701 0.02409 0.01769 0.0128 0.1922 1.0000
12.500 1.3665 0.02536 0.01898 0.0159 0.1803 1.0000
12.750 1.3614 0.02705 0.02069 0.0182 0.1691 1.0000
13.000 1.3531 0.02935 0.02300 0.0198 0.1565 1.0000
13.250 1.3434 0.03218 0.02583 0.0206 0.1454 1.0000
13.500 1.3307 0.03561 0.02926 0.0208 0.1340 1.0000
13.750 1.3201 0.03909 0.03278 0.0207 0.1245 1.0000
14.000 1.3074 0.04292 0.03665 0.0204 0.1164 1.0000
14.250 1.2929 0.04710 0.04086 0.0198 0.1096 1.0000
14.500 1.2802 0.05117 0.04498 0.0191 0.1022 1.0000
14.750 1.2643 0.05570 0.04954 0.0182 0.0963 1.0000
15.000 1.2553 0.05950 0.05343 0.0175 0.0916 1.0000
15.250 1.2412 0.06412 0.05808 0.0164 0.0863 1.0000
15.500 1.2315 0.06832 0.06235 0.0153 0.0811 1.0000
15.750 1.2203 0.07283 0.06691 0.0140 0.0763 1.0000
16.000 1.2115 0.07709 0.07124 0.0128 0.0725 1.0000
16.250 1.2031 0.08140 0.07562 0.0115 0.0678 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 637 AIRFOIL (e637-il)