EPPLER 479 AIRFOIL (e479-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 479 AIRFOIL (e479-il) Reynolds number: 100,000 Max Cl/Cd: 39.02 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e479-il-100000-n5.txt Download as CSV file: xf-e479-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 479 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.250 -0.9791 0.09426 0.08787 -0.0159 1.0000 0.0954
-16.000 -1.0190 0.08402 0.07737 -0.0216 1.0000 0.0978
-15.750 -1.0270 0.07950 0.07279 -0.0236 1.0000 0.1003
-15.500 -1.0274 0.07633 0.06962 -0.0249 1.0000 0.1032
-15.250 -1.0400 0.07132 0.06450 -0.0272 1.0000 0.1061
-15.000 -1.0570 0.06585 0.05885 -0.0298 1.0000 0.1089
-14.750 -1.0672 0.06156 0.05443 -0.0315 1.0000 0.1118
-14.500 -1.0668 0.05888 0.05177 -0.0322 1.0000 0.1148
-14.250 -1.0726 0.05541 0.04821 -0.0334 1.0000 0.1181
-14.000 -1.0810 0.05167 0.04429 -0.0347 1.0000 0.1215
-13.750 -1.0849 0.04862 0.04115 -0.0353 1.0000 0.1248
-13.500 -1.0843 0.04629 0.03882 -0.0356 1.0000 0.1282
-13.250 -1.0865 0.04370 0.03614 -0.0358 1.0000 0.1320
-13.000 -1.0896 0.04110 0.03336 -0.0359 1.0000 0.1360
-12.750 -1.0871 0.03923 0.03150 -0.0354 1.0000 0.1397
-12.500 -1.0856 0.03743 0.02966 -0.0348 1.0000 0.1437
-12.250 -1.0844 0.03566 0.02776 -0.0338 1.0000 0.1483
-12.000 -1.0811 0.03416 0.02619 -0.0326 1.0000 0.1525
-11.750 -1.0768 0.03294 0.02496 -0.0310 1.0000 0.1569
-11.500 -1.0726 0.03175 0.02367 -0.0290 1.0000 0.1620
-11.250 -1.0671 0.03070 0.02255 -0.0269 1.0000 0.1669
-11.000 -1.0602 0.02986 0.02172 -0.0246 1.0000 0.1717
-10.750 -1.0509 0.02897 0.02071 -0.0226 1.0000 0.1778
-10.500 -1.0389 0.02818 0.01991 -0.0208 1.0000 0.1833
-10.250 -1.0260 0.02745 0.01915 -0.0191 1.0000 0.1896
-10.000 -1.0118 0.02670 0.01830 -0.0175 1.0000 0.1962
-9.750 -0.9963 0.02608 0.01771 -0.0160 1.0000 0.2024
-9.500 -0.9804 0.02541 0.01689 -0.0145 1.0000 0.2100
-9.250 -0.9632 0.02487 0.01642 -0.0131 1.0000 0.2161
-9.000 -0.9458 0.02430 0.01575 -0.0117 1.0000 0.2238
-8.750 -0.9276 0.02380 0.01528 -0.0103 1.0000 0.2305
-8.500 -0.9092 0.02330 0.01471 -0.0089 1.0000 0.2382
-8.250 -0.8902 0.02284 0.01425 -0.0076 1.0000 0.2451
-8.000 -0.8712 0.02241 0.01379 -0.0063 1.0000 0.2529
-7.750 -0.8520 0.02198 0.01333 -0.0049 1.0000 0.2601
-7.500 -0.8327 0.02161 0.01295 -0.0036 1.0000 0.2675
-7.250 -0.8135 0.02121 0.01251 -0.0022 1.0000 0.2752
-7.000 -0.7942 0.02089 0.01221 -0.0008 1.0000 0.2823
-6.750 -0.7753 0.02052 0.01176 0.0007 1.0000 0.2902
-6.500 -0.7563 0.02023 0.01154 0.0021 1.0000 0.2969
-6.250 -0.7384 0.01992 0.01116 0.0037 1.0000 0.3049
-6.000 -0.7114 0.01964 0.01096 0.0035 0.9942 0.3120
-5.750 -0.6706 0.01933 0.01053 0.0006 0.9760 0.3224
-5.500 -0.6309 0.01905 0.01032 -0.0018 0.9558 0.3306
-5.250 -0.5908 0.01875 0.00993 -0.0042 0.9342 0.3405
-5.000 -0.5515 0.01849 0.00969 -0.0063 0.9110 0.3489
-4.750 -0.5129 0.01822 0.00933 -0.0083 0.8867 0.3583
-4.500 -0.4787 0.01802 0.00909 -0.0092 0.8602 0.3664
-4.250 -0.4477 0.01782 0.00877 -0.0095 0.8340 0.3749
-4.000 -0.4197 0.01768 0.00858 -0.0092 0.8079 0.3825
-3.750 -0.3930 0.01754 0.00830 -0.0086 0.7835 0.3910
-3.500 -0.3667 0.01742 0.00814 -0.0079 0.7606 0.3981
-3.250 -0.3409 0.01730 0.00790 -0.0073 0.7382 0.4070
-3.000 -0.3150 0.01719 0.00777 -0.0066 0.7173 0.4141
-2.750 -0.2887 0.01709 0.00757 -0.0060 0.6974 0.4231
-2.500 -0.2624 0.01699 0.00744 -0.0054 0.6784 0.4308
-2.250 -0.2362 0.01692 0.00728 -0.0049 0.6604 0.4400
-2.000 -0.2100 0.01684 0.00717 -0.0043 0.6431 0.4483
-1.750 -0.1838 0.01678 0.00704 -0.0038 0.6265 0.4580
-1.500 -0.1576 0.01671 0.00696 -0.0032 0.6109 0.4666
-1.000 -0.1052 0.01663 0.00680 -0.0021 0.5820 0.4862
-0.750 -0.0789 0.01659 0.00673 -0.0016 0.5678 0.4968
-0.500 -0.0526 0.01657 0.00671 -0.0011 0.5544 0.5074
-0.250 -0.0263 0.01655 0.00669 -0.0005 0.5420 0.5182
0.000 0.0000 0.01656 0.00664 0.0000 0.5304 0.5304
0.250 0.0263 0.01655 0.00669 0.0005 0.5182 0.5420
0.500 0.0526 0.01657 0.00671 0.0011 0.5074 0.5544
0.750 0.0789 0.01659 0.00673 0.0016 0.4968 0.5678
1.000 0.1052 0.01663 0.00680 0.0021 0.4862 0.5820
1.500 0.1576 0.01671 0.00696 0.0032 0.4666 0.6109
1.750 0.1838 0.01678 0.00704 0.0038 0.4580 0.6266
2.000 0.2100 0.01684 0.00717 0.0043 0.4483 0.6432
2.250 0.2362 0.01692 0.00728 0.0049 0.4400 0.6604
2.500 0.2624 0.01699 0.00744 0.0054 0.4308 0.6784
2.750 0.2887 0.01709 0.00757 0.0060 0.4231 0.6974
3.000 0.3150 0.01719 0.00777 0.0066 0.4141 0.7173
3.250 0.3409 0.01730 0.00790 0.0073 0.4070 0.7382
3.500 0.3667 0.01742 0.00815 0.0079 0.3981 0.7606
3.750 0.3930 0.01754 0.00830 0.0086 0.3910 0.7835
4.000 0.4197 0.01768 0.00858 0.0092 0.3825 0.8079
4.250 0.4477 0.01782 0.00877 0.0095 0.3749 0.8340
4.500 0.4788 0.01802 0.00909 0.0092 0.3664 0.8603
4.750 0.5129 0.01822 0.00933 0.0083 0.3583 0.8868
5.000 0.5515 0.01849 0.00969 0.0063 0.3489 0.9111
5.250 0.5909 0.01875 0.00993 0.0042 0.3405 0.9342
5.500 0.6309 0.01905 0.01032 0.0018 0.3306 0.9558
5.750 0.6706 0.01933 0.01053 -0.0006 0.3225 0.9760
6.000 0.7114 0.01964 0.01095 -0.0035 0.3120 0.9942
6.250 0.7383 0.01992 0.01116 -0.0037 0.3049 1.0000
6.500 0.7562 0.02023 0.01154 -0.0021 0.2969 1.0000
6.750 0.7752 0.02051 0.01176 -0.0006 0.2902 1.0000
7.000 0.7941 0.02088 0.01221 0.0008 0.2822 1.0000
7.250 0.8135 0.02121 0.01251 0.0022 0.2752 1.0000
7.500 0.8327 0.02161 0.01295 0.0036 0.2675 1.0000
7.750 0.8520 0.02198 0.01334 0.0049 0.2602 1.0000
8.000 0.8713 0.02241 0.01379 0.0063 0.2530 1.0000
8.250 0.8903 0.02284 0.01425 0.0076 0.2452 1.0000
8.500 0.9092 0.02330 0.01472 0.0089 0.2383 1.0000
8.750 0.9276 0.02380 0.01528 0.0103 0.2305 1.0000
9.000 0.9458 0.02429 0.01575 0.0117 0.2238 1.0000
9.250 0.9633 0.02487 0.01642 0.0130 0.2162 1.0000
9.500 0.9805 0.02541 0.01689 0.0145 0.2099 1.0000
9.750 0.9964 0.02608 0.01771 0.0160 0.2024 1.0000
10.000 1.0120 0.02669 0.01830 0.0175 0.1963 1.0000
10.250 1.0262 0.02745 0.01914 0.0191 0.1896 1.0000
10.500 1.0392 0.02818 0.01991 0.0208 0.1833 1.0000
10.750 1.0512 0.02897 0.02071 0.0225 0.1778 1.0000
11.000 1.0606 0.02986 0.02172 0.0245 0.1717 1.0000
11.250 1.0676 0.03070 0.02254 0.0268 0.1668 1.0000
11.500 1.0731 0.03175 0.02367 0.0290 0.1619 1.0000
11.750 1.0773 0.03294 0.02496 0.0309 0.1568 1.0000
12.000 1.0819 0.03415 0.02618 0.0325 0.1525 1.0000
12.250 1.0851 0.03565 0.02775 0.0337 0.1482 1.0000
12.500 1.0863 0.03742 0.02965 0.0347 0.1436 1.0000
12.750 1.0879 0.03921 0.03148 0.0353 0.1396 1.0000
13.000 1.0904 0.04110 0.03336 0.0358 0.1359 1.0000
13.250 1.0871 0.04371 0.03615 0.0357 0.1318 1.0000
13.500 1.0854 0.04626 0.03879 0.0354 0.1281 1.0000
13.750 1.0863 0.04857 0.04109 0.0352 0.1247 1.0000
14.000 1.0821 0.05166 0.04429 0.0345 0.1214 1.0000
14.250 1.0738 0.05538 0.04818 0.0333 0.1179 1.0000
14.500 1.0684 0.05882 0.05170 0.0321 0.1147 1.0000
14.750 1.0693 0.06144 0.05429 0.0314 0.1117 1.0000
15.000 1.0584 0.06585 0.05885 0.0296 0.1088 1.0000
15.250 1.0414 0.07130 0.06448 0.0270 0.1059 1.0000
15.500 1.0297 0.07618 0.06946 0.0248 0.1031 1.0000
15.750 1.0299 0.07926 0.07254 0.0236 0.1002 1.0000
16.000 1.0200 0.08411 0.07746 0.0213 0.0976 1.0000
16.250 0.9806 0.09426 0.08788 0.0157 0.0953 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 479 AIRFOIL (e479-il)