Davis AIRFOIL (davis-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: Davis AIRFOIL (davis-il) Reynolds number: 200,000 Max Cl/Cd: 72.09 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-davis-il-200000-n5.txt Download as CSV file: xf-davis-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: Davis 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4223 0.08519 0.08191 -0.0432 1.0000 0.0148
-9.250 -0.4443 0.08006 0.07689 -0.0433 1.0000 0.0145
-9.000 -0.5393 0.04368 0.03981 -0.0777 0.9814 0.0129
-8.750 -0.5338 0.03811 0.03374 -0.0812 0.9743 0.0130
-8.500 -0.5188 0.03423 0.02945 -0.0833 0.9691 0.0134
-8.250 -0.5017 0.03165 0.02656 -0.0838 0.9632 0.0138
-8.000 -0.4776 0.02930 0.02387 -0.0851 0.9591 0.0143
-7.750 -0.4561 0.02725 0.02149 -0.0853 0.9533 0.0149
-7.500 -0.4293 0.02534 0.01925 -0.0862 0.9490 0.0157
-7.250 -0.4002 0.02392 0.01752 -0.0871 0.9452 0.0173
-7.000 -0.3751 0.02271 0.01598 -0.0869 0.9391 0.0183
-6.750 -0.3457 0.02080 0.01389 -0.0880 0.9359 0.0196
-6.500 -0.3209 0.01982 0.01281 -0.0877 0.9298 0.0210
-6.250 -0.2906 0.01897 0.01183 -0.0885 0.9252 0.0235
-6.000 -0.2570 0.01814 0.01082 -0.0898 0.9222 0.0261
-5.750 -0.2356 0.01713 0.00971 -0.0887 0.9140 0.0282
-5.500 -0.2049 0.01631 0.00879 -0.0895 0.9097 0.0321
-5.250 -0.1788 0.01575 0.00811 -0.0892 0.9024 0.0358
-5.000 -0.1497 0.01501 0.00727 -0.0895 0.8963 0.0403
-4.750 -0.1212 0.01448 0.00666 -0.0897 0.8896 0.0466
-4.500 -0.0934 0.01394 0.00609 -0.0897 0.8822 0.0552
-4.250 -0.0637 0.01338 0.00559 -0.0902 0.8763 0.0775
-4.000 -0.0373 0.01295 0.00528 -0.0900 0.8680 0.1164
-3.750 -0.0065 0.01264 0.00499 -0.0906 0.8619 0.1467
-3.500 0.0206 0.01239 0.00474 -0.0905 0.8535 0.1668
-3.250 0.0507 0.01214 0.00449 -0.0909 0.8468 0.1877
-3.000 0.0785 0.01195 0.00429 -0.0909 0.8380 0.2057
-2.750 0.1072 0.01175 0.00409 -0.0910 0.8292 0.2306
-2.500 0.1368 0.01154 0.00387 -0.0913 0.8199 0.2550
-2.250 0.1638 0.01134 0.00370 -0.0911 0.8087 0.2825
-2.000 0.1914 0.01112 0.00354 -0.0910 0.7987 0.3148
-1.500 0.2460 0.01080 0.00331 -0.0907 0.7799 0.3850
-1.250 0.2729 0.01066 0.00321 -0.0904 0.7702 0.4226
-0.750 0.3247 0.01038 0.00305 -0.0895 0.7501 0.5073
-0.500 0.3487 0.01016 0.00301 -0.0886 0.7406 0.5680
-0.250 0.3702 0.00977 0.00305 -0.0870 0.7313 0.6965
0.000 0.4087 0.00953 0.00317 -0.0886 0.7209 0.8641
0.250 0.4685 0.00957 0.00318 -0.0951 0.7094 0.9451
0.500 0.5174 0.00965 0.00316 -0.0996 0.6967 0.9800
0.750 0.5619 0.00974 0.00316 -0.1032 0.6833 1.0000
1.000 0.5838 0.00984 0.00318 -0.1020 0.6704 1.0000
1.250 0.6055 0.00993 0.00320 -0.1007 0.6568 1.0000
1.500 0.6271 0.01003 0.00324 -0.0995 0.6421 1.0000
1.750 0.6488 0.01013 0.00328 -0.0982 0.6268 1.0000
2.000 0.6704 0.01025 0.00333 -0.0968 0.6104 1.0000
2.250 0.6917 0.01039 0.00338 -0.0955 0.5934 1.0000
2.500 0.7128 0.01055 0.00345 -0.0941 0.5766 1.0000
2.750 0.7339 0.01073 0.00355 -0.0927 0.5610 1.0000
3.000 0.7551 0.01093 0.00367 -0.0913 0.5469 1.0000
3.250 0.7764 0.01114 0.00381 -0.0900 0.5333 1.0000
3.500 0.7977 0.01137 0.00399 -0.0887 0.5199 1.0000
4.000 0.8402 0.01184 0.00437 -0.0861 0.4914 1.0000
4.250 0.8614 0.01208 0.00459 -0.0848 0.4758 1.0000
4.500 0.8825 0.01231 0.00481 -0.0835 0.4593 1.0000
4.750 0.9033 0.01255 0.00504 -0.0822 0.4410 1.0000
5.000 0.9235 0.01281 0.00528 -0.0807 0.4216 1.0000
5.250 0.9429 0.01312 0.00556 -0.0791 0.4028 1.0000
5.500 0.9620 0.01345 0.00584 -0.0775 0.3867 1.0000
5.750 0.9812 0.01380 0.00616 -0.0760 0.3730 1.0000
6.000 1.0004 0.01417 0.00651 -0.0744 0.3604 1.0000
6.250 1.0180 0.01460 0.00691 -0.0726 0.3453 1.0000
6.500 1.0345 0.01509 0.00731 -0.0707 0.3283 1.0000
6.750 1.0506 0.01557 0.00773 -0.0687 0.3098 1.0000
7.000 1.0679 0.01599 0.00815 -0.0669 0.2929 1.0000
7.250 1.0856 0.01638 0.00856 -0.0652 0.2775 1.0000
7.500 1.1017 0.01677 0.00896 -0.0632 0.2585 1.0000
7.750 1.1133 0.01732 0.00943 -0.0604 0.2274 1.0000
8.000 1.1198 0.01819 0.01004 -0.0571 0.1772 1.0000
8.250 1.1083 0.02034 0.01146 -0.0514 0.0782 1.0000
8.500 1.1099 0.02181 0.01268 -0.0478 0.0359 1.0000
8.750 1.1164 0.02296 0.01376 -0.0448 0.0240 1.0000
9.000 1.1263 0.02391 0.01479 -0.0424 0.0209 1.0000
9.250 1.1360 0.02488 0.01588 -0.0401 0.0192 1.0000
9.500 1.1436 0.02601 0.01712 -0.0376 0.0176 1.0000
9.750 1.1488 0.02732 0.01854 -0.0351 0.0164 1.0000
10.000 1.1490 0.02902 0.02039 -0.0322 0.0152 1.0000
10.500 1.1595 0.03190 0.02356 -0.0281 0.0140 1.0000
10.750 1.1616 0.03370 0.02549 -0.0261 0.0133 1.0000
11.000 1.1630 0.03565 0.02756 -0.0242 0.0130 1.0000
11.250 1.1645 0.03769 0.02972 -0.0226 0.0127 1.0000
11.500 1.1648 0.03994 0.03208 -0.0212 0.0123 1.0000
11.750 1.1662 0.04217 0.03441 -0.0200 0.0119 1.0000
12.000 1.1666 0.04458 0.03692 -0.0189 0.0116 1.0000
12.250 1.1676 0.04701 0.03944 -0.0180 0.0113 1.0000
12.500 1.1659 0.04977 0.04226 -0.0173 0.0107 1.0000
12.750 1.1644 0.05264 0.04521 -0.0163 0.0104 1.0000
13.000 1.1661 0.05524 0.04790 -0.0156 0.0100 1.0000
13.250 1.1701 0.05768 0.05048 -0.0151 0.0098 1.0000
13.500 1.1738 0.06021 0.05315 -0.0146 0.0096 1.0000
13.750 1.1771 0.06285 0.05597 -0.0142 0.0094 1.0000
14.000 1.1799 0.06563 0.05890 -0.0139 0.0092 1.0000
14.250 1.1818 0.06860 0.06203 -0.0137 0.0090 1.0000
14.500 1.1825 0.07179 0.06539 -0.0137 0.0088 1.0000
14.750 1.1818 0.07526 0.06903 -0.0139 0.0087 1.0000
15.000 1.1798 0.07897 0.07292 -0.0144 0.0087 1.0000
15.250 1.1754 0.08313 0.07727 -0.0153 0.0085 1.0000
15.500 1.1701 0.08753 0.08186 -0.0164 0.0084 1.0000
15.750 1.1635 0.09220 0.08670 -0.0180 0.0083 1.0000
16.000 1.1558 0.09719 0.09188 -0.0198 0.0082 1.0000
16.250 1.1468 0.10259 0.09747 -0.0220 0.0082 1.0000
16.500 1.1371 0.10827 0.10333 -0.0247 0.0081 1.0000
16.750 1.1298 0.11358 0.10877 -0.0274 0.0079 1.0000
17.000 1.1169 0.12030 0.11569 -0.0308 0.0079 1.0000
17.250 1.1003 0.12818 0.12380 -0.0352 0.0080 1.0000
17.500 1.0960 0.13329 0.12899 -0.0382 0.0078 1.0000
17.750 1.0747 0.14298 0.13893 -0.0440 0.0080 1.0000
18.000 1.0570 0.15219 0.14832 -0.0497 0.0080 1.0000
18.250 1.0334 0.16374 0.16007 -0.0569 0.0083 1.0000
18.500 1.0089 0.17676 0.17323 -0.0648 0.0084 1.0000
|
Polar data table (+)
Polar graphs
<< Back to Davis AIRFOIL (davis-il)