EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Reynolds number: 200,000 Max Cl/Cd: 34.68 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ea61012-il-200000.txt Download as CSV file: xf-ea61012-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER EA 6(-1)-012 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.6744 0.12052 0.11658 0.0005 1.0000 0.0930
-12.250 -0.6527 0.11966 0.11569 0.0028 1.0000 0.0946
-12.000 -0.8951 0.06951 0.06516 -0.0324 1.0000 0.0694
-11.750 -0.9118 0.06485 0.06033 -0.0326 1.0000 0.0670
-11.500 -0.9332 0.06059 0.05592 -0.0318 1.0000 0.0666
-11.250 -0.9565 0.05703 0.05217 -0.0292 1.0000 0.0664
-11.000 -0.9722 0.05326 0.04815 -0.0269 1.0000 0.0663
-10.750 -0.9819 0.04976 0.04433 -0.0246 1.0000 0.0666
-10.500 -0.9859 0.04665 0.04086 -0.0223 1.0000 0.0671
-10.250 -0.9842 0.04403 0.03786 -0.0202 1.0000 0.0676
-9.750 -0.9684 0.03753 0.03076 -0.0174 1.0000 0.0693
-9.500 -0.9499 0.03549 0.02866 -0.0168 1.0000 0.0707
-9.250 -0.9299 0.03383 0.02693 -0.0162 1.0000 0.0718
-9.000 -0.9108 0.03213 0.02507 -0.0154 1.0000 0.0729
-8.750 -0.8908 0.03066 0.02342 -0.0146 1.0000 0.0744
-8.500 -0.8704 0.02920 0.02177 -0.0138 1.0000 0.0758
-8.250 -0.8489 0.02776 0.02013 -0.0130 1.0000 0.0768
-8.000 -0.8266 0.02650 0.01868 -0.0123 1.0000 0.0779
-7.750 -0.8042 0.02565 0.01763 -0.0115 1.0000 0.0790
-7.500 -0.7805 0.02374 0.01566 -0.0112 1.0000 0.0806
-7.250 -0.7567 0.02249 0.01444 -0.0108 1.0000 0.0822
-7.000 -0.7348 0.02158 0.01353 -0.0100 1.0000 0.0836
-6.750 -0.7186 0.02087 0.01282 -0.0083 1.0000 0.0850
-6.500 -0.7032 0.02029 0.01222 -0.0067 0.9977 0.0867
-6.250 -0.6649 0.01949 0.01136 -0.0092 0.9891 0.0894
-6.000 -0.6256 0.01879 0.01057 -0.0117 0.9823 0.0915
-5.750 -0.5908 0.01755 0.00943 -0.0135 0.9746 0.0944
-5.250 -0.5194 0.01621 0.00813 -0.0174 0.9594 0.1025
-5.000 -0.4842 0.01547 0.00742 -0.0192 0.9527 0.1072
-4.750 -0.4544 0.01486 0.00687 -0.0200 0.9432 0.1135
-4.500 -0.4241 0.01430 0.00633 -0.0207 0.9352 0.1225
-4.250 -0.3988 0.01373 0.00588 -0.0205 0.9260 0.1398
-4.000 -0.3798 0.01262 0.00532 -0.0195 0.9176 0.2288
-3.750 -0.3683 0.01126 0.00491 -0.0173 0.9085 0.4206
-3.500 -0.3512 0.01077 0.00509 -0.0150 0.9015 0.5631
-3.250 -0.3273 0.01083 0.00527 -0.0138 0.8944 0.6194
-3.000 -0.3019 0.01100 0.00546 -0.0128 0.8886 0.6522
-2.750 -0.2765 0.01120 0.00567 -0.0119 0.8822 0.6758
-2.500 -0.2505 0.01142 0.00588 -0.0110 0.8764 0.6936
-2.250 -0.2243 0.01167 0.00609 -0.0102 0.8718 0.7094
-2.000 -0.1989 0.01199 0.00642 -0.0092 0.8658 0.7263
-1.750 -0.1742 0.01229 0.00672 -0.0079 0.8591 0.7422
-1.500 -0.1497 0.01256 0.00696 -0.0065 0.8530 0.7566
-1.250 -0.1252 0.01283 0.00729 -0.0051 0.8461 0.7703
-1.000 -0.1012 0.01309 0.00757 -0.0033 0.8405 0.7846
-0.750 -0.0770 0.01325 0.00775 -0.0019 0.8345 0.7960
-0.500 -0.0520 0.01329 0.00776 -0.0013 0.8276 0.8050
-0.250 -0.0260 0.01327 0.00773 -0.0004 0.8222 0.8097
0.000 0.0000 0.01327 0.00776 0.0000 0.8154 0.8154
0.250 0.0259 0.01327 0.00773 0.0004 0.8097 0.8222
0.500 0.0519 0.01329 0.00776 0.0013 0.8050 0.8277
0.750 0.0769 0.01325 0.00775 0.0020 0.7959 0.8345
1.000 0.1011 0.01309 0.00757 0.0034 0.7845 0.8405
1.250 0.1251 0.01284 0.00729 0.0051 0.7705 0.8461
1.500 0.1497 0.01256 0.00696 0.0065 0.7566 0.8530
1.750 0.1742 0.01229 0.00672 0.0079 0.7422 0.8591
2.000 0.1988 0.01198 0.00642 0.0092 0.7262 0.8658
2.250 0.2242 0.01167 0.00609 0.0102 0.7094 0.8718
2.500 0.2504 0.01142 0.00588 0.0111 0.6936 0.8764
2.750 0.2765 0.01120 0.00567 0.0119 0.6758 0.8822
3.000 0.3019 0.01100 0.00546 0.0128 0.6523 0.8886
3.250 0.3273 0.01083 0.00527 0.0138 0.6192 0.8944
3.500 0.3511 0.01077 0.00509 0.0150 0.5628 0.9015
3.750 0.3682 0.01126 0.00491 0.0173 0.4207 0.9085
4.000 0.3799 0.01260 0.00531 0.0195 0.2302 0.9176
4.250 0.3988 0.01373 0.00588 0.0205 0.1399 0.9260
4.500 0.4241 0.01430 0.00633 0.0207 0.1224 0.9353
4.750 0.4544 0.01486 0.00687 0.0200 0.1135 0.9432
5.000 0.4842 0.01547 0.00742 0.0192 0.1072 0.9527
5.250 0.5195 0.01620 0.00813 0.0174 0.1025 0.9594
5.500 0.5534 0.01678 0.00871 0.0159 0.0981 0.9683
5.750 0.5909 0.01754 0.00943 0.0135 0.0944 0.9747
6.000 0.6258 0.01879 0.01057 0.0116 0.0915 0.9824
6.250 0.6650 0.01949 0.01136 0.0091 0.0894 0.9891
6.500 0.7033 0.02028 0.01221 0.0067 0.0867 0.9978
6.750 0.7185 0.02086 0.01282 0.0084 0.0850 1.0000
7.000 0.7348 0.02157 0.01353 0.0100 0.0836 1.0000
7.250 0.7567 0.02248 0.01443 0.0108 0.0822 1.0000
7.500 0.7805 0.02375 0.01566 0.0112 0.0806 1.0000
7.750 0.8041 0.02563 0.01761 0.0115 0.0790 1.0000
8.000 0.8266 0.02649 0.01867 0.0123 0.0778 1.0000
8.250 0.8489 0.02776 0.02013 0.0130 0.0768 1.0000
8.500 0.8704 0.02919 0.02176 0.0138 0.0757 1.0000
8.750 0.8909 0.03064 0.02341 0.0146 0.0743 1.0000
9.000 0.9109 0.03211 0.02504 0.0154 0.0729 1.0000
9.250 0.9300 0.03384 0.02694 0.0162 0.0718 1.0000
9.500 0.9499 0.03549 0.02866 0.0168 0.0707 1.0000
9.750 0.9685 0.03754 0.03076 0.0174 0.0693 1.0000
10.250 0.9841 0.04402 0.03785 0.0202 0.0676 1.0000
10.500 0.9860 0.04667 0.04088 0.0223 0.0671 1.0000
10.750 0.9824 0.04984 0.04441 0.0245 0.0667 1.0000
11.000 0.9723 0.05330 0.04819 0.0268 0.0663 1.0000
11.250 0.9560 0.05696 0.05211 0.0292 0.0663 1.0000
11.500 0.9340 0.06066 0.05599 0.0317 0.0666 1.0000
11.750 0.9130 0.06496 0.06043 0.0325 0.0671 1.0000
12.000 0.8966 0.07018 0.06574 0.0321 0.0677 1.0000
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Polar data table (+)
Polar graphs
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