ONERA OA213 AIRFOIL (oa213-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: ONERA OA213 AIRFOIL (oa213-il) Reynolds number: 100,000 Max Cl/Cd: 41.65 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oa213-il-100000.txt Download as CSV file: xf-oa213-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: ONERA OA213 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3647 0.10170 0.09769 -0.0186 1.0000 0.0885
-8.750 -0.3607 0.09799 0.09401 -0.0225 1.0000 0.0920
-8.500 -0.3734 0.09438 0.09046 -0.0307 1.0000 0.0948
-8.250 -0.2973 0.08747 0.08390 -0.0263 0.9989 0.1008
-8.000 -0.2872 0.08282 0.07925 -0.0299 0.9947 0.1052
-7.750 -0.2969 0.07762 0.07407 -0.0368 0.9887 0.1093
-7.500 -0.3285 0.07255 0.06890 -0.0462 0.9788 0.1108
-7.250 -0.4051 0.08009 0.07634 -0.0319 0.9911 0.1023
-7.000 -0.4025 0.07596 0.07185 -0.0428 0.9811 0.1107
-6.750 -0.3819 0.07002 0.06595 -0.0457 0.9766 0.1139
-6.500 -0.3540 0.06646 0.06240 -0.0479 0.9721 0.1205
-6.250 -0.2946 0.05076 0.04656 -0.0573 0.9437 0.1439
-6.000 -0.3284 0.06021 0.05556 -0.0549 0.9557 0.1446
-5.750 -0.2964 0.05625 0.05180 -0.0565 0.9517 0.1547
-5.500 -0.2844 0.05357 0.04901 -0.0566 0.9428 0.1665
-5.250 -0.2678 0.05111 0.04644 -0.0570 0.9351 0.1825
-5.000 -0.2486 0.04881 0.04406 -0.0573 0.9276 0.2010
-4.500 -0.1888 0.03741 0.03033 -0.0569 0.9127 0.0794
-4.250 -0.1703 0.03506 0.02770 -0.0555 0.9035 0.0758
-4.000 -0.1410 0.03314 0.02536 -0.0552 0.8973 0.0759
-3.750 -0.1184 0.03154 0.02341 -0.0540 0.8883 0.0755
-3.500 -0.0893 0.03003 0.02157 -0.0534 0.8815 0.0758
-3.250 -0.0635 0.02904 0.02033 -0.0525 0.8729 0.0792
-3.000 -0.0362 0.02790 0.01907 -0.0519 0.8655 0.0845
-2.750 -0.0099 0.02706 0.01821 -0.0512 0.8570 0.0912
-2.500 0.0148 0.02626 0.01745 -0.0502 0.8485 0.1018
-2.250 0.0398 0.02544 0.01673 -0.0490 0.8402 0.1177
-2.000 0.0609 0.02476 0.01616 -0.0474 0.8298 0.1443
-1.500 0.2439 0.02145 0.01535 -0.0598 0.8185 1.0000
-1.250 0.2607 0.02130 0.01504 -0.0575 0.8074 1.0000
-1.000 0.2803 0.02125 0.01484 -0.0561 0.7952 1.0000
-0.750 0.2972 0.02106 0.01451 -0.0536 0.7856 1.0000
-0.500 0.3148 0.02084 0.01415 -0.0512 0.7753 1.0000
-0.250 0.3344 0.02072 0.01393 -0.0497 0.7635 1.0000
0.000 0.3490 0.02028 0.01334 -0.0460 0.7565 1.0000
0.250 0.3698 0.02014 0.01313 -0.0447 0.7437 1.0000
0.500 0.3893 0.01993 0.01283 -0.0429 0.7322 1.0000
0.750 0.4051 0.01938 0.01215 -0.0393 0.7249 1.0000
1.000 0.4257 0.01917 0.01187 -0.0378 0.7118 1.0000
1.250 0.4455 0.01888 0.01150 -0.0358 0.6995 1.0000
1.500 0.4641 0.01844 0.01097 -0.0332 0.6889 1.0000
1.750 0.4828 0.01796 0.01039 -0.0305 0.6779 1.0000
2.000 0.5034 0.01765 0.00999 -0.0286 0.6635 1.0000
2.250 0.5240 0.01732 0.00958 -0.0266 0.6486 1.0000
2.500 0.5449 0.01700 0.00917 -0.0247 0.6327 1.0000
2.750 0.5660 0.01669 0.00874 -0.0227 0.6153 1.0000
3.000 0.5878 0.01648 0.00840 -0.0210 0.5952 1.0000
3.250 0.6105 0.01639 0.00820 -0.0197 0.5707 1.0000
3.500 0.6331 0.01636 0.00800 -0.0184 0.5448 1.0000
3.750 0.6559 0.01643 0.00783 -0.0171 0.5173 1.0000
4.000 0.6795 0.01669 0.00788 -0.0164 0.4866 1.0000
4.250 0.7032 0.01706 0.00805 -0.0158 0.4569 1.0000
4.500 0.7270 0.01752 0.00830 -0.0153 0.4304 1.0000
4.750 0.7509 0.01803 0.00857 -0.0148 0.4081 1.0000
5.000 0.7750 0.01861 0.00893 -0.0144 0.3892 1.0000
5.250 0.7994 0.01922 0.00941 -0.0142 0.3720 1.0000
5.500 0.8238 0.01986 0.00988 -0.0139 0.3573 1.0000
5.750 0.8485 0.02050 0.01042 -0.0137 0.3438 1.0000
6.000 0.8734 0.02118 0.01109 -0.0136 0.3316 1.0000
6.250 0.8983 0.02192 0.01175 -0.0134 0.3217 1.0000
6.500 0.9236 0.02262 0.01234 -0.0133 0.3124 1.0000
6.750 0.9482 0.02340 0.01320 -0.0132 0.3033 1.0000
7.000 0.9739 0.02416 0.01380 -0.0131 0.2957 1.0000
7.250 0.9981 0.02501 0.01479 -0.0131 0.2880 1.0000
7.500 1.0235 0.02581 0.01556 -0.0131 0.2817 1.0000
7.750 1.0485 0.02681 0.01659 -0.0131 0.2762 1.0000
8.000 1.0721 0.02774 0.01768 -0.0130 0.2702 1.0000
8.250 1.0977 0.02862 0.01850 -0.0130 0.2651 1.0000
8.500 1.1210 0.02977 0.01978 -0.0130 0.2602 1.0000
8.750 1.1432 0.03085 0.02106 -0.0129 0.2551 1.0000
9.000 1.1672 0.03188 0.02216 -0.0129 0.2511 1.0000
9.250 1.1933 0.03315 0.02337 -0.0130 0.2478 1.0000
9.500 1.2112 0.03469 0.02528 -0.0127 0.2445 1.0000
9.750 1.2296 0.03627 0.02714 -0.0125 0.2412 1.0000
10.000 1.2495 0.03775 0.02881 -0.0123 0.2382 1.0000
10.250 1.2709 0.03912 0.03028 -0.0121 0.2353 1.0000
10.500 1.2948 0.04060 0.03179 -0.0122 0.2327 1.0000
10.750 1.3052 0.04289 0.03442 -0.0116 0.2304 1.0000
11.000 1.3094 0.04554 0.03749 -0.0106 0.2282 1.0000
11.250 1.3103 0.04849 0.04079 -0.0096 0.2264 1.0000
11.500 1.3060 0.05173 0.04434 -0.0084 0.2248 1.0000
11.750 1.2928 0.05550 0.04840 -0.0071 0.2233 1.0000
12.000 1.2145 0.06411 0.05735 -0.0050 0.2235 1.0000
12.250 1.3494 0.05658 0.04946 -0.0070 0.2156 1.0000
12.500 1.3130 0.06183 0.05501 -0.0050 0.2155 1.0000
12.750 1.2677 0.06886 0.06227 -0.0041 0.2162 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA OA213 AIRFOIL (oa213-il)