EPPLER 335 AIRFOIL (e335-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 335 AIRFOIL (e335-il) Reynolds number: 500,000 Max Cl/Cd: 81.33 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e335-il-500000.txt Download as CSV file: xf-e335-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 335 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4302 0.08667 0.08475 -0.0092 1.0000 0.0154
-9.750 -0.4366 0.08096 0.07905 -0.0119 1.0000 0.0154
-9.500 -0.5070 0.08530 0.08323 -0.0118 1.0000 0.0148
-9.250 -0.5194 0.07948 0.07740 -0.0164 1.0000 0.0148
-9.000 -0.5370 0.07414 0.07201 -0.0185 1.0000 0.0149
-8.250 -0.5881 0.06141 0.05891 -0.0115 1.0000 0.0152
-8.000 -0.5979 0.05746 0.05475 -0.0079 1.0000 0.0153
-7.750 -0.6071 0.05319 0.05038 -0.0048 1.0000 0.0155
-7.500 -0.6050 0.05076 0.04792 -0.0025 1.0000 0.0156
-7.250 -0.5997 0.04834 0.04542 -0.0004 0.9938 0.0157
-7.000 -0.5670 0.04520 0.04193 -0.0044 0.8589 0.0163
-6.750 -0.5613 0.04312 0.03950 -0.0012 0.8073 0.0166
-6.500 -0.5533 0.04075 0.03683 0.0018 0.7741 0.0172
-6.000 -0.5387 0.03388 0.02902 0.0098 0.7295 0.0195
-4.750 -0.4426 0.01911 0.01217 0.0215 0.6468 0.0114
-4.500 -0.4129 0.01714 0.00989 0.0217 0.6320 0.0108
-4.250 -0.3840 0.01582 0.00835 0.0219 0.6180 0.0106
-4.000 -0.3574 0.01497 0.00735 0.0225 0.6045 0.0111
-3.750 -0.3324 0.01428 0.00654 0.0232 0.5914 0.0114
-3.500 -0.3077 0.01380 0.00593 0.0241 0.5785 0.0122
-3.250 -0.2858 0.01307 0.00519 0.0253 0.5665 0.0133
-3.000 -0.2621 0.01271 0.00476 0.0263 0.5554 0.0143
-2.750 -0.2388 0.01238 0.00436 0.0272 0.5448 0.0167
-2.500 -0.2150 0.01202 0.00396 0.0282 0.5340 0.0194
-2.250 -0.1921 0.01162 0.00352 0.0294 0.5241 0.0257
-2.000 -0.1732 0.01090 0.00314 0.0312 0.5148 0.1043
-1.750 -0.1609 0.00983 0.00284 0.0339 0.5066 0.3095
-1.500 -0.1769 0.00802 0.00243 0.0425 0.5009 0.6557
-1.250 -0.1547 0.00822 0.00341 0.0453 0.4926 0.8834
-1.000 -0.0567 0.01055 0.00554 0.0332 0.4783 0.9084
-0.750 0.0030 0.01142 0.00624 0.0275 0.4676 0.9164
-0.500 0.0666 0.01213 0.00681 0.0208 0.4573 0.9219
-0.250 0.1026 0.01252 0.00707 0.0196 0.4495 0.9305
0.000 0.1448 0.01256 0.00704 0.0167 0.4414 0.9324
0.250 0.1783 0.01259 0.00695 0.0155 0.4346 0.9343
0.500 0.2100 0.01258 0.00690 0.0147 0.4275 0.9367
0.750 0.2348 0.01265 0.00689 0.0152 0.4217 0.9407
1.000 0.2557 0.01271 0.00690 0.0166 0.4164 0.9449
1.250 0.2908 0.01262 0.00677 0.0150 0.4102 0.9460
1.500 0.3243 0.01259 0.00666 0.0136 0.4044 0.9473
1.750 0.3566 0.01256 0.00660 0.0126 0.3989 0.9488
2.000 0.3874 0.01254 0.00655 0.0118 0.3937 0.9507
2.250 0.4154 0.01259 0.00654 0.0117 0.3887 0.9531
2.500 0.4375 0.01266 0.00661 0.0127 0.3840 0.9562
2.750 0.4598 0.01271 0.00665 0.0137 0.3794 0.9592
3.000 0.4929 0.01264 0.00654 0.0125 0.3746 0.9601
3.250 0.5253 0.01261 0.00649 0.0113 0.3699 0.9613
3.500 0.5571 0.01255 0.00644 0.0103 0.3648 0.9626
3.750 0.5865 0.01255 0.00642 0.0098 0.3602 0.9640
4.000 0.6146 0.01262 0.00646 0.0095 0.3555 0.9656
4.250 0.6417 0.01262 0.00651 0.0095 0.3510 0.9676
4.500 0.6654 0.01268 0.00658 0.0101 0.3465 0.9693
4.750 0.6832 0.01292 0.00679 0.0120 0.3423 0.9721
5.000 0.7153 0.01282 0.00673 0.0108 0.3375 0.9729
5.250 0.7454 0.01276 0.00670 0.0101 0.3324 0.9736
5.500 0.7748 0.01279 0.00672 0.0095 0.3277 0.9745
5.750 0.8046 0.01280 0.00677 0.0088 0.3225 0.9757
6.000 0.8330 0.01281 0.00682 0.0084 0.3173 0.9768
6.250 0.8589 0.01290 0.00691 0.0085 0.3121 0.9778
6.500 0.8846 0.01295 0.00702 0.0087 0.3070 0.9788
6.750 0.9106 0.01300 0.00711 0.0088 0.3009 0.9802
7.000 0.9340 0.01317 0.00727 0.0094 0.2954 0.9814
7.250 0.9571 0.01322 0.00742 0.0101 0.2895 0.9823
7.500 0.9772 0.01340 0.00760 0.0114 0.2832 0.9834
7.750 1.0047 0.01344 0.00771 0.0111 0.2767 0.9843
8.000 1.0335 0.01348 0.00779 0.0105 0.2692 0.9850
8.250 1.0599 0.01356 0.00793 0.0104 0.2615 0.9856
8.500 1.0851 0.01371 0.00809 0.0105 0.2530 0.9861
8.750 1.1104 0.01383 0.00828 0.0106 0.2442 0.9867
9.000 1.1350 0.01403 0.00850 0.0107 0.2346 0.9874
9.250 1.1601 0.01429 0.00877 0.0107 0.2234 0.9884
9.500 1.1833 0.01455 0.00906 0.0111 0.2109 0.9892
9.750 1.2045 0.01489 0.00941 0.0118 0.1973 0.9898
10.000 1.2241 0.01536 0.00984 0.0127 0.1815 0.9905
10.250 1.2428 0.01585 0.01032 0.0137 0.1655 0.9912
10.500 1.2608 0.01644 0.01089 0.0147 0.1493 0.9922
10.750 1.2770 0.01713 0.01155 0.0160 0.1331 0.9934
11.000 1.2913 0.01788 0.01226 0.0175 0.1185 0.9941
11.250 1.3080 0.01869 0.01305 0.0184 0.1043 0.9949
11.500 1.3225 0.01960 0.01394 0.0194 0.0911 0.9958
11.750 1.3353 0.02055 0.01488 0.0207 0.0798 0.9969
12.000 1.3475 0.02152 0.01587 0.0220 0.0704 0.9982
12.250 1.3585 0.02263 0.01701 0.0231 0.0617 0.9997
12.500 1.3461 0.02367 0.01808 0.0285 0.0574 1.0000
12.750 1.3283 0.02463 0.01913 0.0345 0.0552 1.0000
13.000 1.3094 0.02625 0.02081 0.0390 0.0524 1.0000
13.250 1.2907 0.02843 0.02306 0.0423 0.0502 1.0000
13.500 1.2730 0.03090 0.02561 0.0450 0.0480 1.0000
13.750 1.2589 0.03326 0.02805 0.0473 0.0456 1.0000
14.000 1.2412 0.03594 0.03079 0.0496 0.0435 1.0000
14.250 1.2212 0.03886 0.03377 0.0518 0.0422 1.0000
14.500 1.2015 0.04207 0.03704 0.0536 0.0412 1.0000
14.750 1.1856 0.04541 0.04046 0.0544 0.0394 1.0000
15.000 1.1741 0.04869 0.04382 0.0547 0.0371 1.0000
15.250 1.1607 0.05243 0.04761 0.0544 0.0354 1.0000
15.500 1.1444 0.05671 0.05194 0.0538 0.0338 1.0000
15.750 1.1321 0.06077 0.05607 0.0530 0.0324 1.0000
16.000 1.1238 0.06451 0.05989 0.0522 0.0304 1.0000
16.250 1.1133 0.06864 0.06407 0.0511 0.0288 1.0000
16.500 1.1000 0.07326 0.06873 0.0497 0.0275 1.0000
16.750 1.0899 0.07758 0.07313 0.0485 0.0263 1.0000
17.000 1.0819 0.08167 0.07730 0.0472 0.0250 1.0000
17.250 1.0752 0.08565 0.08132 0.0459 0.0234 1.0000
17.500 1.0639 0.09038 0.08611 0.0443 0.0228 1.0000
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