EPPLER 635 AIRFOIL (e635-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 635 AIRFOIL (e635-il) Reynolds number: 500,000 Max Cl/Cd: 88.02 at α=8.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e635-il-500000-n5.txt Download as CSV file: xf-e635-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 635 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5368 0.08341 0.08132 -0.0026 0.9011 0.0066
-9.500 -0.5440 0.07963 0.07717 -0.0052 0.7876 0.0068
-9.250 -0.5759 0.07196 0.06938 -0.0099 0.7625 0.0067
-9.000 -0.5996 0.06717 0.06445 -0.0091 0.7377 0.0066
-8.750 -0.6187 0.06284 0.05996 -0.0072 0.7195 0.0067
-8.500 -0.6317 0.05782 0.05474 -0.0055 0.7030 0.0067
-8.250 -0.6385 0.05309 0.04978 -0.0035 0.6882 0.0067
-8.000 -0.6390 0.04894 0.04540 -0.0015 0.6736 0.0068
-7.750 -0.6333 0.04566 0.04189 0.0004 0.6590 0.0070
-7.500 -0.6288 0.04155 0.03750 0.0029 0.6461 0.0072
-7.250 -0.6168 0.03895 0.03467 0.0047 0.6327 0.0075
-7.000 -0.6187 0.03176 0.02690 0.0092 0.6238 0.0081
-6.750 -0.6254 0.02171 0.01546 0.0159 0.6182 0.0093
-6.500 -0.6042 0.02042 0.01391 0.0170 0.6046 0.0096
-6.250 -0.5805 0.01993 0.01331 0.0178 0.5916 0.0099
-6.000 -0.5565 0.01942 0.01265 0.0185 0.5790 0.0102
-5.500 -0.5074 0.01799 0.01086 0.0200 0.5542 0.0114
-5.250 -0.4820 0.01730 0.00997 0.0206 0.5435 0.0122
-5.000 -0.4562 0.01681 0.00928 0.0212 0.5333 0.0128
-4.750 -0.4315 0.01582 0.00817 0.0219 0.5226 0.0134
-4.500 -0.4060 0.01533 0.00761 0.0224 0.5124 0.0139
-4.250 -0.3805 0.01488 0.00707 0.0229 0.5029 0.0145
-4.000 -0.3548 0.01444 0.00654 0.0235 0.4938 0.0152
-3.750 -0.3289 0.01413 0.00614 0.0240 0.4853 0.0163
-3.500 -0.3030 0.01382 0.00573 0.0246 0.4764 0.0170
-3.250 -0.2780 0.01335 0.00518 0.0253 0.4681 0.0174
-3.000 -0.2542 0.01280 0.00457 0.0262 0.4603 0.0182
-2.750 -0.2293 0.01244 0.00418 0.0269 0.4531 0.0190
-2.250 -0.1784 0.01196 0.00360 0.0281 0.4387 0.0211
-2.000 -0.1530 0.01175 0.00331 0.0287 0.4313 0.0219
-1.750 -0.1271 0.01156 0.00307 0.0292 0.4249 0.0225
-1.500 -0.1013 0.01138 0.00283 0.0298 0.4186 0.0233
-1.250 -0.0755 0.01121 0.00261 0.0303 0.4129 0.0249
-1.000 -0.0494 0.01107 0.00245 0.0308 0.4065 0.0275
-0.750 -0.0232 0.01099 0.00231 0.0312 0.4006 0.0308
-0.500 0.0024 0.01079 0.00220 0.0318 0.3954 0.0527
-0.250 0.0269 0.01053 0.00211 0.0324 0.3901 0.1154
0.000 0.0395 0.00932 0.00192 0.0350 0.3857 0.4287
0.250 0.0784 0.00772 0.00207 0.0329 0.3799 0.8962
0.500 0.1292 0.00820 0.00249 0.0286 0.3736 0.9324
0.750 0.1635 0.00859 0.00282 0.0278 0.3689 0.9453
1.000 0.2102 0.00890 0.00304 0.0240 0.3629 0.9492
1.250 0.2464 0.00911 0.00316 0.0224 0.3578 0.9533
1.500 0.2723 0.00917 0.00318 0.0228 0.3539 0.9566
1.750 0.2986 0.00923 0.00320 0.0232 0.3497 0.9593
2.000 0.3308 0.00931 0.00322 0.0222 0.3449 0.9603
2.250 0.3623 0.00939 0.00324 0.0214 0.3406 0.9613
2.500 0.3936 0.00945 0.00328 0.0206 0.3365 0.9626
2.750 0.4241 0.00952 0.00333 0.0200 0.3325 0.9641
3.000 0.4533 0.00961 0.00338 0.0196 0.3286 0.9659
3.250 0.4811 0.00970 0.00345 0.0196 0.3248 0.9681
3.500 0.5033 0.00980 0.00355 0.0208 0.3210 0.9715
3.750 0.5350 0.00987 0.00360 0.0198 0.3169 0.9722
4.000 0.5665 0.00997 0.00366 0.0189 0.3133 0.9731
4.250 0.5977 0.01004 0.00375 0.0181 0.3099 0.9741
4.500 0.6281 0.01012 0.00383 0.0174 0.3058 0.9752
4.750 0.6580 0.01021 0.00392 0.0168 0.3016 0.9765
5.000 0.6874 0.01035 0.00405 0.0163 0.2980 0.9780
5.250 0.7161 0.01043 0.00417 0.0160 0.2945 0.9796
5.500 0.7434 0.01054 0.00430 0.0159 0.2904 0.9814
5.750 0.7681 0.01070 0.00447 0.0164 0.2862 0.9835
6.000 0.7991 0.01081 0.00458 0.0156 0.2817 0.9843
6.250 0.8299 0.01089 0.00470 0.0147 0.2762 0.9851
6.500 0.8598 0.01103 0.00485 0.0140 0.2707 0.9860
6.750 0.8896 0.01115 0.00500 0.0133 0.2659 0.9870
7.000 0.9192 0.01128 0.00516 0.0127 0.2598 0.9881
7.250 0.9482 0.01146 0.00535 0.0121 0.2541 0.9894
7.500 0.9770 0.01160 0.00555 0.0116 0.2483 0.9906
7.750 1.0046 0.01181 0.00577 0.0113 0.2417 0.9919
8.000 1.0322 0.01198 0.00600 0.0110 0.2350 0.9931
8.250 1.0595 0.01224 0.00626 0.0107 0.2275 0.9943
8.500 1.0892 0.01243 0.00651 0.0099 0.2185 0.9951
8.750 1.1178 0.01270 0.00679 0.0092 0.2090 0.9961
9.000 1.1455 0.01302 0.00711 0.0087 0.1985 0.9971
9.250 1.1726 0.01341 0.00750 0.0082 0.1842 0.9981
9.500 1.1989 0.01396 0.00799 0.0077 0.1627 0.9992
9.750 1.2228 0.01464 0.00859 0.0075 0.1420 1.0000
10.000 1.2382 0.01549 0.00934 0.0089 0.1207 1.0000
10.250 1.2539 0.01626 0.01006 0.0103 0.1052 1.0000
10.500 1.2680 0.01709 0.01085 0.0118 0.0900 1.0000
10.750 1.2808 0.01797 0.01170 0.0136 0.0768 1.0000
11.000 1.2926 0.01882 0.01255 0.0154 0.0667 1.0000
11.250 1.3016 0.01980 0.01350 0.0176 0.0562 1.0000
11.500 1.3091 0.02078 0.01447 0.0200 0.0491 1.0000
11.750 1.3141 0.02180 0.01551 0.0226 0.0423 1.0000
12.000 1.3155 0.02288 0.01663 0.0257 0.0379 1.0000
12.250 1.3053 0.02396 0.01777 0.0303 0.0351 1.0000
12.500 1.2924 0.02584 0.01971 0.0334 0.0323 1.0000
12.750 1.2865 0.02809 0.02205 0.0345 0.0301 1.0000
13.000 1.2813 0.03072 0.02478 0.0349 0.0277 1.0000
13.250 1.2705 0.03417 0.02827 0.0348 0.0241 1.0000
13.500 1.2658 0.03713 0.03133 0.0346 0.0232 1.0000
13.750 1.2571 0.04059 0.03487 0.0343 0.0212 1.0000
14.000 1.2458 0.04441 0.03876 0.0337 0.0194 1.0000
14.250 1.2340 0.04835 0.04278 0.0331 0.0176 1.0000
14.500 1.2180 0.05286 0.04733 0.0322 0.0146 1.0000
14.750 1.2067 0.05687 0.05142 0.0314 0.0134 1.0000
15.000 1.1972 0.06075 0.05539 0.0306 0.0125 1.0000
15.250 1.1847 0.06519 0.05989 0.0294 0.0108 1.0000
15.500 1.1719 0.06983 0.06458 0.0281 0.0088 1.0000
15.750 1.1628 0.07407 0.06890 0.0269 0.0080 1.0000
16.000 1.1537 0.07840 0.07330 0.0256 0.0073 1.0000
16.250 1.1451 0.08272 0.07768 0.0242 0.0065 1.0000
16.500 1.1329 0.08766 0.08269 0.0226 0.0055 1.0000
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