EPPLER 399 AIRFOIL (e399-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 399 AIRFOIL (e399-il) Reynolds number: 50,000 Max Cl/Cd: 16.93 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e399-il-50000-n5.txt Download as CSV file: xf-e399-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 399 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.3086 0.12540 0.11867 -0.0544 1.0000 0.0554
-11.250 -0.3201 0.12208 0.11547 -0.0540 1.0000 0.0559
-11.000 -0.3328 0.11897 0.11248 -0.0533 1.0000 0.0562
-10.750 -0.3361 0.11817 0.11177 -0.0508 1.0000 0.0577
-10.500 -0.3457 0.11648 0.11018 -0.0488 1.0000 0.0586
-10.250 -0.3468 0.11275 0.10650 -0.0505 0.9971 0.0597
-10.000 -0.3408 0.10702 0.10078 -0.0557 0.9908 0.0616
-9.750 -0.3409 0.10027 0.09406 -0.0617 0.9838 0.0632
-9.500 -0.3507 0.09206 0.08590 -0.0683 0.9751 0.0641
-9.250 -0.3696 0.08226 0.07616 -0.0761 0.9654 0.0645
-9.000 -0.3943 0.07287 0.06676 -0.0843 0.9535 0.0641
-8.750 -0.4222 0.06381 0.05753 -0.0940 0.9389 0.0634
-8.500 -0.4365 0.05544 0.04867 -0.1035 0.9263 0.0640
-8.250 -0.4288 0.05174 0.04474 -0.1065 0.9165 0.0672
-8.000 -0.4146 0.04850 0.04117 -0.1094 0.9086 0.0722
-7.750 -0.3974 0.04540 0.03771 -0.1119 0.9008 0.0777
-7.500 -0.3716 0.04228 0.03402 -0.1155 0.8953 0.0860
-7.250 -0.3533 0.04133 0.03312 -0.1153 0.8869 0.0929
-7.000 -0.3226 0.03934 0.03084 -0.1179 0.8822 0.1031
-6.750 -0.3006 0.03775 0.02892 -0.1187 0.8750 0.1132
-6.500 -0.2731 0.03648 0.02739 -0.1200 0.8689 0.1252
-6.250 -0.2380 0.03525 0.02597 -0.1224 0.8650 0.1398
-6.000 -0.2188 0.03469 0.02540 -0.1218 0.8572 0.1517
-5.750 -0.1888 0.03385 0.02441 -0.1231 0.8518 0.1679
-5.500 -0.1522 0.03295 0.02330 -0.1254 0.8483 0.1876
-5.250 -0.1335 0.03264 0.02290 -0.1246 0.8403 0.2032
-5.000 -0.1036 0.03223 0.02246 -0.1255 0.8352 0.2221
-4.750 -0.0675 0.03178 0.02192 -0.1273 0.8317 0.2443
-4.500 -0.0498 0.03176 0.02183 -0.1262 0.8239 0.2616
-4.250 -0.0196 0.03149 0.02139 -0.1270 0.8188 0.2844
-4.000 0.0156 0.03123 0.02107 -0.1284 0.8153 0.3066
-3.750 0.0323 0.03138 0.02117 -0.1270 0.8078 0.3241
-3.500 0.0598 0.03133 0.02104 -0.1272 0.8025 0.3449
-3.250 0.0952 0.03113 0.02067 -0.1287 0.7990 0.3688
-3.000 0.1132 0.03137 0.02084 -0.1275 0.7920 0.3869
-2.750 0.1382 0.03146 0.02088 -0.1273 0.7864 0.4062
-2.500 0.1710 0.03136 0.02072 -0.1281 0.7829 0.4272
-2.250 0.1914 0.03163 0.02094 -0.1272 0.7767 0.4454
-2.000 0.2125 0.03188 0.02114 -0.1264 0.7705 0.4637
-1.750 0.2445 0.03186 0.02104 -0.1271 0.7668 0.4851
-1.500 0.2807 0.03173 0.02081 -0.1284 0.7641 0.5085
-1.250 0.2842 0.03256 0.02167 -0.1252 0.7546 0.5223
-1.000 0.3154 0.03258 0.02163 -0.1258 0.7507 0.5447
-0.750 0.3492 0.03250 0.02153 -0.1265 0.7480 0.5666
-0.500 0.3520 0.03346 0.02251 -0.1234 0.7387 0.5825
-0.250 0.3803 0.03355 0.02259 -0.1234 0.7346 0.6053
0.000 0.4141 0.03349 0.02251 -0.1242 0.7318 0.6311
0.250 0.4149 0.03458 0.02367 -0.1208 0.7225 0.6491
0.500 0.4402 0.03472 0.02384 -0.1202 0.7183 0.6744
0.750 0.4712 0.03462 0.02376 -0.1203 0.7155 0.7035
1.000 0.4696 0.03590 0.02513 -0.1167 0.7060 0.7269
1.250 0.4918 0.03602 0.02531 -0.1155 0.7017 0.7598
1.500 0.5184 0.03581 0.02517 -0.1146 0.6989 0.8002
1.750 0.5106 0.03723 0.02672 -0.1102 0.6888 0.8435
2.000 0.5360 0.03704 0.02661 -0.1096 0.6845 1.0000
2.250 0.5782 0.03710 0.02649 -0.1124 0.6821 1.0000
2.500 0.5798 0.03914 0.02847 -0.1110 0.6713 1.0000
2.750 0.6138 0.03945 0.02865 -0.1124 0.6677 1.0000
3.000 0.6416 0.04006 0.02917 -0.1130 0.6632 1.0000
3.250 0.6487 0.04178 0.03087 -0.1119 0.6540 1.0000
3.500 0.6826 0.04199 0.03100 -0.1129 0.6506 1.0000
4.000 0.7136 0.04450 0.03346 -0.1117 0.6361 1.0000
4.250 0.7487 0.04456 0.03350 -0.1126 0.6332 1.0000
4.500 0.7462 0.04693 0.03589 -0.1107 0.6221 1.0000
4.750 0.7774 0.04719 0.03615 -0.1111 0.6182 1.0000
5.250 0.8069 0.04982 0.03882 -0.1097 0.6033 1.0000
5.500 0.8421 0.04974 0.03878 -0.1103 0.6003 1.0000
5.750 0.8376 0.05238 0.04145 -0.1084 0.5883 1.0000
6.000 0.8707 0.05238 0.04151 -0.1088 0.5847 1.0000
6.250 0.8688 0.05490 0.04409 -0.1071 0.5733 1.0000
6.500 0.9000 0.05498 0.04423 -0.1073 0.5691 1.0000
7.000 0.9305 0.05746 0.04688 -0.1058 0.5532 1.0000
7.500 0.9619 0.05982 0.04941 -0.1044 0.5371 1.0000
8.000 0.9946 0.06199 0.05180 -0.1029 0.5207 1.0000
8.250 0.9947 0.06452 0.05441 -0.1017 0.5085 1.0000
8.500 1.0284 0.06391 0.05394 -0.1014 0.5041 1.0000
8.750 1.0276 0.06657 0.05671 -0.1003 0.4913 1.0000
9.000 1.0642 0.06547 0.05577 -0.0998 0.4874 1.0000
9.250 1.0624 0.06825 0.05865 -0.0987 0.4739 1.0000
9.750 1.1000 0.06937 0.06008 -0.0970 0.4567 1.0000
10.000 1.1010 0.07191 0.06274 -0.0960 0.4433 1.0000
10.500 1.1407 0.07241 0.06357 -0.0941 0.4255 1.0000
11.000 1.1521 0.07634 0.06779 -0.0923 0.3999 1.0000
11.250 1.1883 0.07406 0.06571 -0.0908 0.3936 1.0000
11.750 1.2014 0.07762 0.06958 -0.0890 0.3665 1.0000
12.000 1.2171 0.07806 0.07018 -0.0879 0.3543 1.0000
12.250 1.2543 0.07525 0.06754 -0.0861 0.3444 1.0000
12.500 1.2702 0.07555 0.06799 -0.0849 0.3299 1.0000
12.750 1.2836 0.07622 0.06877 -0.0837 0.3142 1.0000
13.000 1.2937 0.07740 0.07002 -0.0827 0.2975 1.0000
13.250 1.2989 0.07941 0.07210 -0.0820 0.2802 1.0000
13.500 1.3038 0.08148 0.07425 -0.0813 0.2625 1.0000
13.750 1.3087 0.08363 0.07642 -0.0807 0.2448 1.0000
14.000 1.3125 0.08597 0.07875 -0.0802 0.2273 1.0000
14.250 1.3147 0.08862 0.08138 -0.0798 0.2106 1.0000
14.500 1.3152 0.09162 0.08436 -0.0797 0.1946 1.0000
14.750 1.3143 0.09495 0.08768 -0.0798 0.1797 1.0000
15.000 1.3124 0.09854 0.09127 -0.0802 0.1657 1.0000
15.250 1.3102 0.10228 0.09504 -0.0807 0.1528 1.0000
15.500 1.3083 0.10604 0.09881 -0.0813 0.1409 1.0000
15.750 1.3069 0.10974 0.10249 -0.0820 0.1299 1.0000
16.000 1.3070 0.11314 0.10581 -0.0827 0.1199 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 399 AIRFOIL (e399-il)