ONERA OA212 AIRFOIL (oa212-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: ONERA OA212 AIRFOIL (oa212-il) Reynolds number: 50,000 Max Cl/Cd: 26.21 at α=10.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oa212-il-50000.txt Download as CSV file: xf-oa212-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: ONERA OA212 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5001 0.12568 0.11905 0.0033 1.0000 0.2337
-9.500 -0.4568 0.11857 0.11188 0.0060 1.0000 0.2434
-9.250 -0.4877 0.11818 0.11162 0.0008 1.0000 0.2520
-9.000 -0.4473 0.11198 0.10535 0.0034 1.0000 0.2659
-8.750 -0.4422 0.10825 0.10166 0.0021 1.0000 0.2759
-8.500 -0.4654 0.10719 0.10071 -0.0015 1.0000 0.2889
-8.250 -0.4283 0.10169 0.09516 0.0005 1.0000 0.3030
-8.000 -0.4186 0.09799 0.09148 0.0000 1.0000 0.3161
-7.750 -0.4141 0.09477 0.08831 -0.0005 1.0000 0.3326
-7.500 -0.4040 0.09161 0.08518 -0.0002 1.0000 0.3531
-7.250 -0.3977 0.08885 0.08246 0.0003 1.0000 0.3771
-7.000 -0.4202 0.08758 0.08131 0.0007 1.0000 0.4008
-6.750 -0.3749 0.08292 0.07661 0.0028 1.0000 0.4310
-6.500 -0.3626 0.08016 0.07387 0.0048 1.0000 0.4638
-6.250 -0.3492 0.07803 0.07177 0.0078 1.0000 0.5081
-6.000 -0.3266 0.07576 0.06951 0.0114 1.0000 0.5645
-5.000 -0.5166 0.05323 0.04551 -0.0125 1.0000 0.1956
-4.750 -0.5038 0.04872 0.04050 -0.0123 1.0000 0.1813
-4.500 -0.4879 0.04598 0.03729 -0.0116 1.0000 0.1777
-4.250 -0.4709 0.04327 0.03426 -0.0109 1.0000 0.1770
-4.000 -0.4519 0.04072 0.03135 -0.0103 1.0000 0.1762
-3.750 -0.4310 0.03866 0.02877 -0.0097 1.0000 0.1781
-3.500 -0.4101 0.03657 0.02653 -0.0093 1.0000 0.1823
-3.250 -0.3877 0.03499 0.02476 -0.0090 1.0000 0.1871
-3.000 -0.3638 0.03370 0.02310 -0.0086 1.0000 0.1947
-2.750 -0.3400 0.03230 0.02155 -0.0084 1.0000 0.2022
-2.500 -0.3163 0.03138 0.02041 -0.0081 1.0000 0.2133
-2.250 -0.2928 0.03037 0.01940 -0.0078 1.0000 0.2252
-2.000 -0.2689 0.02959 0.01857 -0.0075 1.0000 0.2401
-1.750 -0.2450 0.02899 0.01799 -0.0071 1.0000 0.2590
-1.500 -0.2208 0.02850 0.01754 -0.0067 1.0000 0.2829
-1.250 -0.1970 0.02798 0.01718 -0.0064 1.0000 0.3150
-1.000 -0.1179 0.02555 0.01761 -0.0110 1.0000 1.0000
-0.750 -0.1066 0.02589 0.01750 -0.0091 1.0000 1.0000
-0.500 -0.0807 0.02646 0.01771 -0.0101 0.9956 1.0000
-0.250 -0.0202 0.02756 0.01842 -0.0175 0.9794 1.0000
0.000 0.0385 0.02870 0.01928 -0.0244 0.9626 1.0000
0.250 0.0887 0.02971 0.02009 -0.0295 0.9439 1.0000
0.500 0.1381 0.03069 0.02093 -0.0344 0.9249 1.0000
0.750 0.1831 0.03156 0.02170 -0.0382 0.9048 1.0000
1.000 0.2398 0.03239 0.02244 -0.0437 0.8855 1.0000
1.250 0.2732 0.03303 0.02303 -0.0451 0.8638 1.0000
1.500 0.3156 0.03357 0.02354 -0.0477 0.8422 1.0000
1.750 0.3601 0.03397 0.02392 -0.0503 0.8215 1.0000
2.000 0.4014 0.03419 0.02413 -0.0519 0.8014 1.0000
2.250 0.4384 0.03428 0.02422 -0.0525 0.7822 1.0000
2.500 0.4677 0.03438 0.02432 -0.0518 0.7630 1.0000
2.750 0.4847 0.03479 0.02473 -0.0497 0.7414 1.0000
3.000 0.5089 0.03496 0.02489 -0.0482 0.7229 1.0000
3.250 0.5338 0.03501 0.02496 -0.0467 0.7059 1.0000
3.500 0.5641 0.03466 0.02461 -0.0452 0.6917 1.0000
3.750 0.5888 0.03458 0.02454 -0.0434 0.6766 1.0000
4.000 0.6078 0.03494 0.02491 -0.0415 0.6603 1.0000
4.250 0.6259 0.03542 0.02542 -0.0396 0.6445 1.0000
4.500 0.6433 0.03606 0.02608 -0.0379 0.6294 1.0000
4.750 0.6629 0.03658 0.02662 -0.0362 0.6154 1.0000
5.000 0.6997 0.03549 0.02555 -0.0346 0.6066 1.0000
5.250 0.7137 0.03667 0.02678 -0.0332 0.5916 1.0000
5.500 0.7276 0.03793 0.02810 -0.0318 0.5772 1.0000
5.750 0.7456 0.03889 0.02914 -0.0305 0.5647 1.0000
6.000 0.7828 0.03782 0.02808 -0.0289 0.5554 1.0000
6.250 0.7917 0.03985 0.03021 -0.0280 0.5411 1.0000
6.500 0.8028 0.04171 0.03216 -0.0269 0.5281 1.0000
6.750 0.8464 0.04005 0.03053 -0.0253 0.5192 1.0000
7.000 0.8553 0.04220 0.03281 -0.0244 0.5056 1.0000
7.250 0.8583 0.04502 0.03574 -0.0236 0.4920 1.0000
7.500 0.8802 0.04582 0.03663 -0.0223 0.4803 1.0000
7.750 0.9328 0.04314 0.03399 -0.0205 0.4688 1.0000
8.000 0.9307 0.04646 0.03749 -0.0195 0.4544 1.0000
8.250 0.9282 0.04982 0.04098 -0.0185 0.4405 1.0000
8.500 0.9510 0.05027 0.04153 -0.0169 0.4259 1.0000
8.750 0.9854 0.04926 0.04063 -0.0149 0.4090 1.0000
9.000 1.0236 0.04764 0.03905 -0.0127 0.3888 1.0000
9.250 1.0796 0.04391 0.03508 -0.0104 0.3627 1.0000
9.500 1.0903 0.04490 0.03622 -0.0083 0.3399 1.0000
9.750 1.1263 0.04313 0.03417 -0.0063 0.3106 1.0000
10.000 1.1342 0.04402 0.03520 -0.0039 0.2840 1.0000
10.250 1.1486 0.04418 0.03524 -0.0016 0.2547 1.0000
10.500 1.1630 0.04437 0.03519 0.0008 0.2265 1.0000
10.750 1.1674 0.04588 0.03671 0.0031 0.2047 1.0000
11.000 1.1819 0.04686 0.03743 0.0049 0.1852 1.0000
11.250 1.1866 0.04910 0.03974 0.0067 0.1718 1.0000
11.500 1.1857 0.05199 0.04282 0.0086 0.1622 1.0000
11.750 1.1973 0.05429 0.04507 0.0098 0.1527 1.0000
12.000 1.1843 0.05781 0.04892 0.0119 0.1483 1.0000
12.250 1.2018 0.06000 0.05099 0.0127 0.1403 1.0000
12.500 1.1773 0.06448 0.05585 0.0144 0.1389 1.0000
12.750 1.1486 0.06988 0.06156 0.0149 0.1383 1.0000
13.000 1.1132 0.07666 0.06862 0.0138 0.1388 1.0000
13.250 1.0725 0.08518 0.07735 0.0110 0.1401 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA OA212 AIRFOIL (oa212-il)